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Part No. PWA 182408 PWA OPER: INSTR. 200 



Gas “Turbine Engine 



June, 1952 

Rewritten FEBRUARY 1958 

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go genie PLES a ee eg ix 

The Flight Spectrum .............0....0.:c cect eet tenn l 
The Basic Gas Turbine Engine ....................---.eiee eee 4 
How The Gas Turbine Operates .........0000.........::ceeeeceetete tees 5 
eee ee rrr re ere rere tern eet rear 7 
Types Of Gas Turbine Engines .....................0: res 11 
Turboprop and Turboshaft Engine Nose Sections 15 
a eis ivi Es MAR Ge 16 
CS = ie ee 19 
Se Oe eT ere eerie ee 26 
Fuel Manifolds and Nozzles ...........:....500:0c0 eects 26 
NI ne eg cs st a eee 27 
a 9 pe a a 30 
ee i, 8 = See) ego 33 
Se UE oo SS, sas Snes HOR I 36 
A I nS ee ease pe ae 38 
Sa we acs co 12 Wao  ss g 4] 
I eo ae asec eR ERs 45 
i Sage econ ree TE eer eer ges Sere 46 
Fuel Systems and Fuel Controls ...................::::st rs 48 
Turboprop Fuel Controls and Propeller Governors .................. 58 
Lubrication Systems ................0.....5.00: citer eee eeereeree tree eens 59 
Famition Systems ee eerie ener 61 
Engine Coolitig «.............-...:ccccrece ie tenercceeerteaeteenerter sere 64 
Engine Insulation Blankets ...................00:: 0s 66 
Water or Coolant Injection ........................ es 66 
Protection Against Icing ..................:::::ecccc etter 69 
Turboprop Asymmetric Thrust Control Devices ....................... 70 
Turboprop Engine Brakes ....................:c cette Pie S 



SECTION Ili — GAS TURBINE ENGINE PERFORMANCE..._—sds_izi(i(y(yyy((N(. 73 
IID ci, 2:7 0c) ce OIE ret le ch 73 
ee 80 
Gas Turbine POMOmMNee oie cocececceeccc 82 
Basic Engine Analysis — Turbojet 2.000000. 84 
Basic Engine Analysis — Turboprop .............................. 92 
Component Performance Details 0. 101 
SECTION IV — GAS TURBINE ENGINE—si—isi(is(i(y#y(;(«w¥téjw#t “ 109 
I cot care een ore ee et tet ee 110 
Engine Operating Variable (Dual Axial Compressor Engines) 112 
Emgume Tnsisumentotion. «oo... 114 
i A ee 116 
gap Patines NR RRG ETE ot ne (li 21 RMR ober x du ea 117 
I occ. eee et eee ar 118 
Letdown, Approach and ES SE ag ey ee ee era 120 
See ee eee ae A Fe BE 121] 
Emgiee Tellet Amiiteine oi ea. 121 
Trimming Dual Axial Compressor Engines... 123 
Cieaming Engine Air Passat noo... 130 
RPM for Checking Thrust on Dual Axial Compressor Engines 131 
Exhaust Gas Temperature Sense coccccccccccccsecscseceee, 132 
III I 33233, ces es eos 134 
SGrErOr TamrntereS oes 136 
We RU Ne noo cc S. 139 
MS nk Re 140 
I 5550555 5s. uid eg eee ee eke 142 
es oasis es ak cc 143 
EIN ooo criss a cone aa a ee 145 



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cK Fe Fina a ies iene reer ean 
1-2. Typical Turbojet Engine Internal Pressure WTI ios os: 7 
1—3. Single Centrifugal or Axial Compressor Turboprop or Turbojet 

Station Designations .....................::cccccceeeereeee teense eee e eee neeeeneenenees 11 
1—4. Dual Compressor Turboprop or Turbojet Station Designations... 11 
1-5. Centrifugal Compressor Turbojet .................::0:: cece 12 
1-6. Single Axial Compressor Turbojet .............0..0.. ces 12 
1—7. Dual Axial Compressor Turbojet ..................::: cette 12 
1-8. Dual Axial Compressor Turbojet with Afterburner ....................., 12 
1-9. Single Axial Compressor, Direct Propeller Drive Turboprop ........ 13 
1—10. Dual Axial Compressor, Direct Propeller Drive Turboprop ........ 13 
1-11. Single Axial Compressor, Free-turbine Propeller Drive Turboprop 13 
1-12. Turbofan with Mixed Exhaust ..........0...0..0...:ccc eee 14 
1—13. Turbofan with Nonmixed Exhaust ....................-eeee es 14 
1-14. Bypass Gas Turbine Engine ...................00:: cece 14 
2-1. Engine Nose Section and Propeller Reduction Gearing, P&WA 

Turboprop Engine (T34) ............: ccc eet tes 15 
2-2. Single Entrance Air Inlet Duct ............0..0.. ccs 16 
2-3. Wing Root Divided Entrance Duct ...............0... es 17 
2-4. Fuselage Side-scoop Divided Entrance Duct .............0 es 7 
2-5. Bellmouth Compressor Inlet and Screen for Ground Test Stand 18 
2—6. Typical Turboprop Ducted Spinner and Air Inlet Configuration 18 
2-7. Streamline or Conical Spinner and Air Inlet Configuration ........ 18 
2-8. Underscoop Air Inlet Used With Offset Reduction Gear Turboprop 18 
2-9. Major Components of a Single-entry Centrifugal Compressor ...... 20 
2-10. Double-entry Centrifugal Compressor Impeller ......................... 20. 
2—11. Multistage Centrifugal Compressor ...................:::cc es 20 
2—12. Components and Assembly of Axial-flow Compressor .................. 21 
2—13. Axial-flow Compressor Cross Section .............:.::::ccceseees 21 
2-14. Airfoil Cross Section of Axial Compressor Rotor Blade .............. 22 
2-15. Dual Axial Compressor System ................:::::ceccsssseeee ee ee eee e tet teteen 23 
2—16. Cross-sectional Area Comparison of Single and Dual Axial Com- 

pressor EMgines o.oo... cece cence tees ee ccereeecseceeneee cena niente ete 24 
2-17. Typical, Multiple-can-type Combustion Chamber Assembly ........ 28 
2-18. Individual Can-type Burner ..............0.. enti 28 
2-19. Typical Annular Combustion Chamber 2... es 29 


List of DLustrations 












Cross Section of Annular Combustion Chamber... 29 
Typical Can-annular Combustion Chamber Assembly... 29 
Individual Can-annular-type Burner 30 
I nn ere Pe a eed in ake 30 
Impulse Turbine Rotor Blades 2... occ. 31 
Reaction Turbine Rotor Blades ...000000200o00oooooooooooooooocceeccececee. 31 
Two-part Turbine for Dual Compressor Engine... 32 
Shrouded Turbine Rotor Blades 00000000000... oooooococcocecceceeeecee. 32 
Conventional Convergent Exhaust Duct 0... 34 
Convergent-divergent Exhaust Duct (Nozzle)... 34 
Suet Mevetser in Gee |... cclcccccce 36 
W-Clamshell Mechanical Blockage Thrust Reverser .............. 38 
Noise Level, with Average Background ............................. 39 
Turbojet Exhaust Noise Pattern eco. 39 
Puene Lvl V5 Tomei Tie oc cee. 40 
Two-position Adjustable Afterburner Nozzle... 43 
Axial Compressor Accessory Drive Airbleed Ports ................ 45 
Axial Compressor Mechanical Accessory Drive... ........... 46 
Typical Starting Sequence for a Gas Turbine Engine... 47 
Fuel System, Nonafterburning Engine Equipped For Water In- 
SN ce ES. Coenen one |. ee ew 49 
Fuel System, Afterburning Engine ..00.0...0.00o0o0ooooooocooccceceee. 51 
Hamilton Standard JFC 12-11 Fuel Control 56 
matee Pel Comirot Wa i 
Typical Lubrication System For a Dual Axial Compressor Engine 60 
Typical Lubrication System Breather and Pressurizing Valve 60 
Typical High-energy, Capacitor-type Ignition System 62 
ree I cn a re 64 
Constramed-gap Igniter Plag 00... ooo. 64 
Typical Outer Case Temperatures For a Dual Axial Compressor 
Te cenit ee at Se toe es 65 
Typical Engine Nacelle Cooling Arrangement... 66 
Typical Engine Insulation Blanket |... 67 
Typical System for Water Injection at the Compressor Inlet and the 
INE oye CAL ee ee ent ae 68 
Compressor Inlet and Diffuser Case Water Injection Regulators 69 
Pe See 3 re ee ee ee 8 71 
pene we a I co. 72 
pe Se ee 74 
Comparison of Thermometer Scales ........0... occ 74 
Pressure Ratio in a Reciprocating Engine... 75 














Temperature and Pressure Rise Due to the Effect of Ram .......... 76 
Engine Operating Cycles ..........0....:..::cccccee eerie 78 
Variation of Wing or Compressor Blade Efficiency with Reynolds 
a regs hc To -au sata Temes Th ins CART TRY EE 79 
Variation of Speed of Sound with Temperature ......................5. 79 
Compression and Expansion Curve for Gas Turbine Engines ...... 82 
Turbojet Station Designations for Amalysis.................0...0c 84 
Corrected Turbojet Performance Curves for a Hypothetical Dual 
Sitar Caer TA oon. en eres 89 
I gS rere ore enter ret tery eer tres 90 
Engine Airflow Parameter ......................:0:ccsceec etter etnies 90 
I a ca a. pica ca sv ergs ex renice vaee RT pd bot ee 90 
Estimated Engine Corrections for Airbleed and Power Extraction 91 
Estimated Thrust Correction Factor Due to Duct Loss ................ 92 
Turboprop Station Designations for Analysis .....................0005 92 
Ambient Temperature Effect on Shaft Horsepower, Net Thrust 
sg a aS Alpe tee a SORA, amt ie rete re Setereer sr rte a: Tey Tey 97 
Basic Turboprop Engine Curves ......................c ce eteeer errr 98 
Turboprop Specification Curve .................:.:cec eres 99 
Turboprop Shaft Horsepower Correction Factors ......................... 99 
Shaft Horsepower Correction for Engine Pressure Ratio .............. 99 
Typical Turboprop Power Setting Curves .................0 es 100 
in ree oe pt ree PETES Tra eT 103 
I a aca cage occas cheng Prot ee 104 
Temperature Conversion Table ......................:: ccc eects 105 
Ideal Ram Pressure Ratio vs True Airspeed ......................:5005. 106 
Compressor Inlet Temperature vs True Airspeed ..................0.0. 107 
Typical Thrust Setting Curves For Engine Pressure Ratio .............. 118 
Aircraft Climb Curve (Military Engine Ratings) ........................ 118 
Typical Temperature-RPM Curve ..................0: eerie 126 
Typical Take-off Thrust Check Curve ..................: ieee tee 127 
Turbine Blade Damage Caused by Excessive Temperatures .......... 133 
Comparative Net Thrust at Sea Level ..................0-:: ccs 137 
Comparative Thrust Specific Fuel Consumption .....................0.4. 137 
Reciprocating and Turboprop Engine Operating Ranges .............. 138 
Typical P&kWA PT2 or T34 Turboprop Performance .................... 138 
Propeller Blade Angle Variation ....................0.00.. sen 139 
Taree Taner TOWER ID. oe in ee ees 140 










The Aircraft Gas Turbine Engine and Its Operation, Pratt & Whitney Aircraft 
Operating Instructions 200, has been rewritten to provide aircraft flight and 
ground crews with a practical basic manual for aircraft gas turbine engines. 
Numerous good texts are available to the student of theory and design of air- 
craft gas turbines, on the one hand, and to operational crews in the form of 
Pratt & Whitney Aircraft Engine Maintenance Manuals, General and Specific 
Operating Instructions, aircraft Flight Handbooks and military Techanical 
Orders, on the other. The Aircraft Gas Turbine Engine and Its Operation is 
intended to fill the gap between the two. Since basic aircraft gas turbine en- 
gine designs are now fairly well established, this book brings previously pub- 
lished material up to date and covers the subject in somewhat greater detail 
than has been done before. 

A comprehensive knowledge of how and why an engine operates is a funda- 
mental prerequisite of intelligent, efficient and safe flying. The same axiom 
also applies to engine maintenance. It is hoped that this book will serve to 
fulfill this requirement. 


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—100 —80 —60—40 -—-20 0 20 40 60 80 100 120 150 200 250 300 400500 oO 

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PWA OPER. | NSTR. 200 



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The steady progress of powered flight has closely followed the develop- 
ment of suitable aircraft powerplants. Unlike the eternal question of the hen 
and the egg, there is no doubt as 
to which was necessary first. 
Without a lightweight and yet 
adequately powerful engine, 
controlled flight of sufficient dis- 
tance to serve a useful purpose 
would not be possible. For lack 
of an adequate means of propulsion, the machine conceived by Leonardo da 
Vinci could not have flown, even if it had been otherwise capable. Although 
Germany’s Dr. N. A. Otto created the four-stroke internal combustion engine 
in 1876, it was not until twenty years later that Daimler was able to perfect the 
eight-horsepower engine which 
enabled the Wolfert “Deutsch- 
land” to make the first gasoline- 
powered dirigible flight. Wilbur 
and Orville Wright had to devel- 
op their own engine before they 
could achieve successful flight at 
Kitty Hawk in 1903. Later, Glenn 
H. Curtiss met with outstanding 
success due largely to the engines 
which he was instrumental in de- 
veloping. And so it has gone, 
down through the pages of aviation history. Larger and more efficient engines 
lead to larger, faster and higher flying aircraft. 

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The Flight Spectrum 

The pros and cons of powerplant types for aircraft have been hotly debated 
since the earliest days of powered flight. The reciprocating engine, turboprop, 
turbojet, ramjet and the rocket, each has its limitations as well as uses for 
which it is best suited. The Flight Spectrum (Figure 1-1) compares the prob- 


. JNSTR. 200 


able ultimate potential of various types of airborne vehicles. Maximum speed 
and altitude capabilities shown for individual aircraft may not necessarily be 
achieved simultaneously or in continuous level flight. 

The reciprocating engine, which may have reached its ultimate in size and 
horsepower, will long be with us as the workhorse of low and medium alti- 
tudes and airspeeds. Along with the shaft turbine, it is finding new usefulness 
in helicopters and some types of convertiplanes. The turboprop combines the 
advantage, inherent in propeller-driven aircraft, of short take-offs with the 
higher and faster flying capability of the gas turbine engine. The turbojet, 
with its increased efficiency at quite high altitudes and airspeeds, is ideal for 
high-flying, high-performance military aircraft and fast, long-range commer- 
cial airliners. A ramjet engine is particularly suited to high altitude and high 
speed but, first, must be carried aloft by some means other than its own thrust 
to reach a velocity sufficient to enable the engine to get started and operate. 
The uses and applications of rocket engines are well known. They vary from 
take-off assists for aircraft normally propelled by conventional powerplants 
to reaction engines for supplying the propulsive forces necessary to carry 
earth satellites to their operational altitude and speed. The future possibilities 
of the rocket are practically limitless. 

Man is a creature who lives miles deep on the bottom of an ocean of air 
which forms a protective canopy over the surface of the earth. Raise him a 
few miles above the bottom of his ocean and he cannot survive unless some 
means are provided to duplicate, approximately, the air temperature and 
pressure of his normal environment. First, man devised wagons, ships, trains 
and automotive vehicles to transport himself and his wares along the bottom 
of his ocean of air. Then he developed the dirigible and the airplane, which 
would travel in this ocean itself. All of the means of propulsion which enable 
people to travel need air or oxygen, whether it be men progressing on foot 
or in an animal-drawn vehicle or the engines powering boats, trains or ramjet 
aircraft. All must have air to provide a means of oxidizing (burning) their 
fuels to obtain energy. Above the altitude limitations of the human body, the 
vehicle must supply pressurized oxygen or air for its passengers and crew. 
Above the air limitations of the engine which propels it, the vehicle must 
carry all of its fuel and air (or other means of supporting combustion) with 
it, as in the case of the rocket. 



> Aircraft or missiles can be operated in continuous level flight only in a 
restricted area of the altitude-flight speed spectrum. The minimum speed 
boundary of this level flight “corridor” is reached when the combined effect 
of wing lift and centrifugal force is no longer sufficient to support aircraft 
weight. Transient flight is possible at lower flight speeds by use of a ballistic- 
type flight path, where altitude is being varied throughout the flight, or by 
aircraft supported directly by powerplant thrust. Except at very high altitudes, 
the maximum speed for continuous flight occurs where the increase in air- 
craft and powerplant structure weight required to overcome the adverse 
effects of high ram air pressure and temperature becomes excessive. The 
effects of pressure predominate at low altitudes, whereas the rapid deteriora- 
tion of the strength of structural materials at high temperatures is the primary 
factor at high altitudes. Development of better materials and improved con- 
struction techniques will tend to raise these maximum speed limits. At 
very high altitudes, the maximum speed for continuous level flight is limited 
to the orbiting velocity. This is the speed at which centrifugal force just 

r counterbalances the earth’s gravitational attraction. Transient operation is 


a. | a a ee a a he 

possible at speeds above orbiting velocity. A vehicle reaching escape velocity 
in a horizontal or ascending trajectory will have sufficient energy to leave the 

earth’s gravitational field. 


Variations in ambient temperature with increasing altitude are an impor- 

en tant factor in flight. As the Flight Spectrum shows, temperature does not vary 

uniformly over a wide range of altitude. In the standard atmosphere, the air 

becomes steadily colder up to an altitude of about 36,000 feet and then re- 

mains at a constant 69.7°F below zero up to about 82,000 feet, after which 

it begins to rise in temperature. The warming trend continues until plus 

49.1°F is reached at a little over 154,000 feet. From this altitude up to 

almost 174,000 feet, the air temperature again remains constant. It then once 

| more becomes colder until 103.7°F below zero is reached at an altitude of 

about 245,000 feet. This temperature holds for, roughly, 35,000 more feet. 

P Somewhere not far above 280,000 feet, the air starts to get hot again. Not 

| too much is known about the data with respect to temperatures at altitudes 

r above 300,000 feet. Although the temperatures at altitudes of 200 miles, 

= Te 

and more, get very hot, the low molecular concentration of the rarefied air 

has a greatly reduced heating effect. Temperature at these very high altitudes 

is something quite different from that with which we are familiar near sea 

level. Were it not for this fact, earth satellites and space rockets would not 
- be possible. 

High-speed aircraft are usually operated at predetermined Mach numbers 

P instead of specific airspeeds. Mach number is the ratio of flight speed to the 
| speed of sound (Mach 1.0). Sonic velocity and Mach number vary with air 
- temperature. Therefore, at standard day conditions, the airspeed which cor- 
eo responds to a given Mach number will vary with the air temperature as alti- 


PWA OPER. | NSTR. 200 

tude is gained. Airspeed is measured in knots which are nautical miles per 

hour. (1.0 knots equals 1.15 statute miles per hour. ) 

A helicopter (shown on the Flight Spectrum) is defined as an aircraft which 
obtains lift by means of a rotor alone. A compound helicopter obtains lift by 
the combination of a rotor and a wing. Airspeeds and altitudes for other 
types of “combination aircraft,” which may be powered either by recipro- 
cating, turboprop or turbojet engines, are not shown because the capabilities 
of each type will vary widely, depending upon the individual aircraft design 
characteristics. Vertical take-off and landing (VTOL) aircraft, short take-off 
and landing (STOL) aircraft, and convertiplanes are in this category. 

The Basic Gas Turbine Engine 

The term, “gas turbine,” 
could be misleading. At 
first thought, because the 
word, “gas,” is so often 
used for gasoline, there 
could be the impression 
that the reference is to a 
gasoline turbine engine. 
The name, however, 
means exactly what it says; that is, a turbine type of engine which is operated 
by a gas, differentiated, for instance, from one operated by steam vapor or 
water. The gas which operates the turbine usually is the product of the com- 
bustion which takes places when a suitable fuel is burned with the air passing 
through the engine. In most gas turbines, the fuel is not gasoline at all, but, 
rather, a low-grade distillate such as military JP-4 or commercial kerosene. 

Both the reciprocating 
engine and the gas tur- 
bine develop power or 
thrust by burning a 
combustible mixture of 
fuel and air. Both con- 
vert the energy of the 
expanding gases into 
propulsive force. The 
reciprocating engine does this by changing the energy of combustion into 
mechanical energy which is used to turn a propeller. Aircraft propulsion is 
obtained as the propeller imparts a relatively small amount of acceleration 
to a large mass of air. The gas turbine, in its basic turbojet configuration, im- 
parts a relatively large amount of acceleration to a smaller mass of air, and 
thus produces thrust or propulsive force directly. Here, the similarity between 
the two types of engine ceases. 

The reciprocating engine is a complicated machine when compared with 
the gas turbine, which, if only the basic mechanically coupled compressor 
and turbine is considered, may be thought of as having only one major mov- 
ing part. (There would be two moving parts if it were a dual axial compres- 


a aT LZ 





sor engine.) Air comes in through a hole in the front of the engine and goes 
out, greatly heated and accelerated, through a hole in the rear. Somewhere 
between the two holes, the engine develops thrust which causes it to try to 
move forward. To give a better understanding of how this is accomplished, 
the engine might be considered as consisting, literally, of hundreds of holes 
from front to rear. For instance, each passage between the blades of the 
compressor or those of the turbine is essentially a hole within the engine. By 
applying Newton’s Second and Third Laws and the F = Ma equation, it is 
fairly easy to understand how and why the engine develops its over-all thrust. 
It is much more difficult to understand exactly what happens at each point 
inside the engine as the air, fuel and gases pass through each series of the 
many holes. One must possess more than a casual acquaintance with thermo- 
dynamics, aerodynamics and the laws pertaining to the physics of gases to 
be able to comprehend completely the production of mechanical energy in a 
gas turbine. 

How The Gas Turbine Operates 

Basically, a gas turbine engine may 
be considered as consisting of four 
main sections: a compressor, a 
burner, a turbine and a tailpipe 
having a jet nozzle. Turbojet ver- 
sions of the gas turbine engine are 
devices to generate pressures and gases and thereby provide mass and accel- 
eration. Turbojet engines derive their propulsive force through the application 
of Sir Isaac Newton’s Third Law which states that for every action there is 
an equal and opposite reaction. When applied to all gas turbine engines, 
this can be expressed by saying that for every force generated there will be 
an equal and opposite force. The opposite force, in the turbojet engine, is en- 
gine thrust. If the gas turbine is a turboprop or turboshaft, the turbine drives a 
propeller or a drive shaft as well as the compressor, as will be shown later. 

It is generally believed that Sanford A. Moss was responsible for building 
the first gas turbine in the United States. In May of 1918, Dr. Moss super- 
vised the construction by the General Electric Company of a turbosuper- 
charger for reciprocating engines, driven by an 
exhaust gas turbine. This is considered the earli- 
est application of a gas turbine for aircraft use. 
The turbosupercharger and the turbojet engine 
are closely related, since they both convert hot 
gases into mechanical work by means of a tur- 
bine. The main difference is that the turbojet 
generates its own heat and converts it into 
thrust, whereas the turbosupercharger depends 
upon the hot gases from the exhaust of a recip- 
rocating engine. The turbosupercharger may 
be so installed that some jet thrust is developed 



INSTR. 200 

as the gases are exhausted. In both types, the expansion of the hot gases 
through the turbine drives the compressor to compress more air. The remain- 
ing energy is used to produce a high-velocity jet stream through a nozzle. 

Newton’s Second Law states that a change in motion is proportional to the 
force applied. Expressed as an equation, force equals mass multiplied by 
acceleration (F = Ma). Force is the net thrust. Acceleration is a term that 
means the rate of change of veloc- 
ity. The velocity change is be- 
tween the low velocity of the in- 
coming air, the zero velocity of 
the fuel and the high velocity of 
the outgoing gases, all velocities 
being relative to that of the en- 
2 gine. When velocity changes are 
substituted in the equation in in place of acceleration, the idea of momentum 
changes within the engine being equal to force or thrust can be understood. 

Mass, in the case of the turbojet, is the mass of air plus the mass of fuel 
which pass through the engine. Acceleration of these masses is accomplished 
in two ways. First, the air mass is compressed and pressure is built up as the 
air goes through the compressors with little change in velocity. Secondly, the 
fuel and part of the air are burned to produce heat. The heated gases expand 
in the burner section and accelerate through the turbine inlet nozzle at the 
outlet of the burner section. The turbines extract power to drive the compres- 
sors. This decelerates the gases but leaves some pressure. The jet nozzle 
allows the gases to attain their final acceleration and generates the out- 
going momentum. 

The incoming momentum of the air and the zero momentum of the fuel 
entering the engine must be subtracted from the outgoing momentum of the 
gases in order to arrive at the over-all change in momentum which represents 
thrust. The thrust developed by a turbojet engine, then, may be said to re- 
sult from the unbalanced forces and momentums created within the engine 
itself. When the static pressure in the throat of the jet nozzle exceeds the 
ambient outside air pressure, an additional thrust is developed at this point. 
Figure 1-2 graphically represents the manner in which the internal pressures 
vary throughout the engine. These pressures and the areas on which they 
work are indicative of the momentum changes within the engine. Since 
engine pressure is proportional to engine thrust, Figure ]-2 indicates how 
the over-all thrust produced by the engine is developed. The final unbalance 
of these pressures and areas gives, as a net result, the total thrust which 
the engine is developing. In practice, this unbalance may be measured or 
calculated in terms of pressure to enable the pilot to monitor engine thrust. 




Ss Ga 


Turboprop engines func- 
tion in a similar manner, 
the chief difference being 
that the jet thrust produced 
is held to a minimum. 
Their relatively large tur- 
bines are designed to ex- 
tract all of the power pos- 
| sible from the expanding 
7 gases flowing from the 
burner section. This power | 
is used to rotate the pro- = 
peller which, in turn, accel- 
erates a large mass of air 
to produce thrust to propel : 
the aircraft. Figure 1—2. Typical Turbojet Engine Internal 
Pressure Variations 

= _ BAe 


ool eal 

cig An rena a as 
- — 
a ee 

The Terms Used 

To be able to understand a subject, one must, first, be familiar with the 
terms used. As in other fields, gas turbine engines have a language of their 
own. Special meanings are applied to common words. The following para- 
graphs will serve to explain some of these words and terms without involv- 
ing perplexing mathematical theory. Terms applying to the physics of gases 
will be discussed in Section III. 

rea < > Stare 
— ieee _—_ 


SET re ce ane as 


One speaks of horsepower when describing a reciprocating engine or a turbo- 
prop. Power is defined as work per unit of time and involves a distance. 
Expressed as an equation, this means 

ae a 

where: P = Power 

F = Foree 
D = Distance 
T = Time 

One horsepower is the unit used to describe the equivalent of 33,000 foot- 
pounds of work performed in one minute, or 550 foot-pounds of work in one 
second. In a reciprocating engine or turboprop, it is possible to measure dis- 
tance and time. Torque and rpm are used in computing horsepower. However, 
these same distance and time elements make use of the terms, “power” and 
“horsepower,” unacceptable for a turbojet engine. When a turbojet engine is 
static, as in the case of an aircraft parked on the ground or when an engine 
is mounted in a ground test stand, distance and time are zero because no 
movement is involved which can be measured against a period of time. 
Although torque and rpm are produced by the turbine, the horsepower de- 
veloped is used entirely within the engine itself. According to the definition 



and equation for power, none, as such, is being produced; yet, a forward force 
is being exerted when the engine is operating. It might be said that thrust is 
the measurement of the amount that an engine pushes against its attachment 
points. The propulsive force developed by a turbojet is measured in pounds 
of thrust. | 

There are two kinds of thrust: net thrust, and gross thrust. Net thrust is 
the thrust which results from the change in momentum of the mass of air 
and the mass of fuel which pass through the engine plus an additional force 
at the jet nozzle represented by the difference in static pressure at the noz- 
zle and the ambient static pressure. The change in momentum which gen- 
erates thrust is expressed as the exhaust or outgoing momentum minus the 
total incoming momentum of the air and fuel. Force or thrust equals mass 
multiplied by acceleration, and acceleration is represented by the difference 
between the outgoing and the incoming velocities. So, by substituting in 
the F = Ma equation, an equation can be written to show how this part of 
the net thrust of a turbojet engine is developed. 

= Se Exhaust _ Incoming Air , Incoming Fuel 
f= Gas Momentum Momentum ~ Momentum 
where: F, = Net thrust in lbs 

Since mass is a function of weight and gravitational acceleration, momentum 
equals the weight of a mass divided by gravitational acceleration (g) and 
multiplied by velocity. The outgoing exhaust gases will have the mass or 
weight of both the air and the fuel consumed by the engine. The incoming air 
velocity is approximately equivalent to the speed of the aircraft. The incoming 
fuel velocity is considered zero because the fuel is carried aboard the aircraft 
and, therefore, will have no initial velocity. The equation then becomes: 

i, a @- 7 @+0 

or, transposing: 
R= (W-W) + (Y) 
where: w, = Airflow through the engine in los per sec 
w; = Fuel flow in lbs per see 
§ = 32.2 (gravitational acceleration in ft per see per sec) 

V; = Exhaust gas velocity in ft per see 
V, = Aircraft velocity (airspeed) in ft per see 



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—_ = aa ca 


In most practical cases, particularly when conventional convergent jet noz- 
zles are used, all of the pressure within the engine cannot be converted to 
velocity. The unconverted pressure is the pressure which represents the 
additional thrust mentioned earlier. This pressure must be added to that 
produced by the changes in momentum. The additional pressure and the 
thrust which it generates become more pronounced as airspeed is gained, 
attaining real significance at ‘the higher Mach numbers. So, to complete 
the story, the net thrust equation must be written: | 

F, = + (Vj—V,) + 2 (V,) + A; (P; — Pn) 

where: A; = Area of jet nozzle in sq ft 
P, = Static pressure at the jet nozzle discharge in lbs per sq ft 
Pin= Foe (ambient) air static pressure at the nozzle in Ibs per sq ft 

It is common practice not to consider fuel flow when computing thrust 
because, for all practical purposes, the weight of air leakage through the 
engine can be assumed to be approximately equivalent to the weight of the 
fuel consumed. Therefore, the net thrust equation now becomes: 

Gross thrust is the thrust developed at the engine exhaust nozzle. This in- 
cludes both the thrust generated by the outgoing momentum of the exhaust 
gases and the additional force resulting from the difference between the 
static pressure at the nozzle and the static pressure of the ambient air. Gross 
thrust does not take the incoming momentum of the air and fuel into con- 
sideration. Zero incoming momentum is assumed, which is true only when 
the engine is static. Without considering fuel flow, the equation for gross 
thrust is: 

F,= ? (V;) + A; (P, — Pen) 
where: Fy = Gross thrust in lbs 

When the aircraft and engine are static, net thrust and gross thrust are 
equal. When the term, “thrust,” is used by itself in discussing a gas turbine 
engine, the reference is usually to net thrust, unless otherwise stated. 

Once the aircraft and engine are in motion, time and distance enter the 
picture, and an approximate comparison can be made with the horsepower 
which a reciprocating engine develops. The equivalent of 33,000 foot-pounds 




per minute, or one horsepower, is 375 mile-pounds per hour. The standard 
power equation can be written to show that, at an airspeed of 375 miles 
per hour, one pound of thrust equals 
one horsepower. This is called thrust 
horsepower. Similarly, at an airspeed 
of 750 miles per hour, one pound of 
thrust equals two thrust horsepower, 
and so on. 


@375 MPH 1” Fn=1 THP 
@750 MPH 1* Fn=2 THP 

Static engine thrust is measured directly in an engine test stand. Stands are 
usually constructed in such a manner that they float, pushing against a 
calibrated scale which accurately measures the thrust in pounds. Thrust 
stands are also available to measure the static thrust exerted by a complete 
aircraft and engine installation, although they are little used because of the 
complications involved. Once an installed engine becomes airborne, direct 
measurement of thrust is no longer practical. Consequently, as will be ex- 
plained later, compressor rpm and turbine discharge pressure (or engine 
pressure ratio), which are proportional to the thrust being developed, are 

instrumented and used to indicate the propulsive force which an engine © 

is producing. 

Thrust Specific Fuel Consumption 

To enable an accurate comparison to be made between turbojet engines, 
fuel consumption is reduced to a common denominator, applicable to all 
types and sizes of turbojet engines. The term used is thrust specific fuel 
consumption. This is the engine fuel flow divided by the net thrust. The 
result, known as TSFC, is the amount of fuel required to produce one pound 
of thrust. One method of comparing engines is on the basis of their TSFC 
at various thrust settings and flight conditions. 

ISFC = Fr. 

where: TSFC = Thrust specific fuel consumption 

w; = Fuel flow in los per hour 
F, = Net thrust in lbs 

Engine Station Designations 

Numerical station designations are assigned to facilitate specific reference 
to the various sections of a gas turbine engine. These coincide with the 
location of the various engine components, as shown in Figures 1-3 and 1-4. 



_—— = = = = = = = = 


—~- «mm wep TD 

—— Ga 


—_ = = = ww 



The static pressure at the com- 
pressor inlet, for instance, is Ps2, 
and the total pressure at this point 
is Piz. Similarly, the total temper- 
ature at the turbine discharge of 
a dual axial compressor engine is 
Tiz. Compressor rpm is referred 
to as “N,” or, in the case of a dual 
axial compressor engine, as Ni 
and Ne for the low pressure and am 
high pressure compressors, re- 

Inlet Diffuser & Duct, 
ny Reduction Gear 

"Exhaust Duct 
op Exhaust Duc 

Compressor Burner 
2 4 

spectively. Station designations Figure 1—3. Single Centrifugal or Axial 
for single axial compressor and Compressor Turboprop or Turbojet 
centrifugal compressor turbo- Station Designations 

props and turbojets are the same. 
For single axial compressor en- 
gines with afterburners, Station 
5 is the entrance to the basic ex- 
haust duct, Station 6 is the en- 
trance to the afterburner combus- 
tion chamber, Station 7 is the 

Inlet Diffuser 
And Duct 

entrance to the jet nozzle and posndooracrey 
: : : ee pressor Burner 
Station 8 is the plane of the jet 
nozzle. For dual axial compressor 
afterburner engines, these same Figure 1—4. Dual Compressor Turbo- 
stations are numbered 7 through prop or Turbojet Station Designations 

10, respectively. Ahead of the en- 

gine are station designations for the ambient or outside air, designated by an 
inferior “am,” and for the entrance to the aircraft inlet air duct, designated by 
an inferior “1.” Another commonly used term, not usually associated with a 
station designation, is “Pp,” meaning the engine internal pressure in the en- 
gine burner section. 

Gas Generator 

The basic engine (which excludes an inlet duct and jet nozzle or propeller 
shaft and gearing) is sometimes called a “gas generator” because it is this 
portion of the engine which generates the gas to be used for propulsive 
thrust or power. 

Types Of Gas Turbine Engines 

If a gas turbine engine relies entirely upon jet thrust to develop its propul- 
sive force, it is known as a turbojet. These engines are, in turn, further 

classified by the type of compressor which they employ. The centrifugal 


. JNSTR. 200 




Jet Exhaust 


compressor (Figure 1-5) works very 
well in the smaller turbojet and tur- 
boprop engines where a high com- 
pression ratio is not too essential. 
This design was standard for early 
aircraft gas turbines. Large, high- 
performance engines require the 
greater efficiency and higher com- 
‘pression ratios attainable only with 
an axial-flow type of compressor. 
Axial compressors have the added 
advantages of light weight and small frontal area. Either a single compressor 
(Figure 1-6) or a dual compressor (Figure [-7) may be used. The latter type 
results in higher compressor efficiencies, compression ratios and thrusts. In 
dual compressor engines, one tur- 
bine or set of turbine wheels drives 
the high pressure compressor, and 
another the low pressure compres- 
sor, both rotor systems operating 
independently of one another ex- 
cept for air flow. The turbine for 
the low pressure compressor is the 
rear turbine and is connected to its 
compressor by a shaft passing 
through the hollow center of the 
high pressure compressor and tur- Figure 1—7. Dual Axial 
bine assembly drive shaft. Compressor Turbojet 


Figure 1—5. Centrifugal Compressor 

Figure 1—6. Single Axial 
Compressor Turbojet 

Frequently, a turbojet engine is equipped with an afterburner for increased 
thrust (Figure 1-8). This increase in thrust can be accomplished regardless 
of the type of compressor used. Roughly, about 25 per cent of the air enter- 
ing the compressor and passing through the engine is required for combus- 
tion. Only this amount of air is required to attain the maximum temperature 
which can be tolerated by the metal parts. The balance of the air is needed 
primarily for cooling purposes. Essentially, an afterburner is simply a huge 
Stovepipe, attached to the rear of the engine, through which all of the 

Figure 1—8. Dual Axial Compressor Turbojet with Afterburner 






exhaust gases must pass. Fuel is injected into the forward section of the 

—_—ae wee a ES Ts 

afterburner and is ignited. Combustion is possible because 75 per cent of the 
air which originally entered the engine still remains unburned. The result 
is, in effect, a tremendous blowtorch which increases the total thrust pro- 
duced by the engine by approximately 50 per cent or more. Although the 
total fuel consumption increases almost 2/2 times, the net result is profitable 
for special bursts of aircraft speed, climb or acceleration. A turbojet air- 
craft with an afterburner can reach a given altitude with the use of less fuel 
by climbing rapidly in “afterburning” than by climbing much more slowly 
in “nonafterburning.” The weight and noise of an afterburner, which, on 
long flights, is used only occasionally, precludes the device being employed 
in present-day, transport-type aircraft. 
When the turbine shaft is coupled to a propeller as well as to the compressor, 

f the gas turbine engine becomes a turboprop. This conversion can be accom- 
plished with either a single or multistage centrifugal compressor, a single axial 

f compressor or a dual axial compressor. In most cases, the propeller reduction 
and drive gearing is connected 
directly to the compressor drive 

f shaft (Figure 1-9) or, when a 
dual axial compressor is used, 

P to the low pressure compressor 

drive shaft (Figure 1-10). On 

eo still another type, the propeller 

f is driven independently of the Figure 1—9. Single Axial Compressor, Direct 
compressor by a free turbine of Propeller Drive Turboprop 

P its own (Figure 1-11). If a tur- 

bine of a gas turbine engine is 
connected to a drive shaft 

P which, in addition to the com- 

, pressor, drives something other 

r than a propeller, such as the 

| rotors of a helicopter or a Figure 1—10. Dual Axial Compressor, Direct 
ground air-compressor unit, the Propeller Drive Turboprop 

P engine is referred to as a shaft 

turbine or turboshaft engine. 

F In principle, the turbofan ver- ai UU EEE 

| sion of a gas turbine engine is 9\_|WP 

P the same as the turboprop ex- 

cept that the ratio of secondary 
airflow to primary combustion Figure 1—11. Single Axial Compressor, 

P airflow is much lower and the Free-turbine Propeller Drive Turboprop 

| geared propeller is replaced by 

- 4 duct-enclosed, axial-flow fan driven at engine speed. In such an engine, only 

| en 30 to 60 per cent of the available propulsive energy is diverted to the fan, so 


ea Vy 


the propulsive efficiency and thrust specific fuel consumption fall between 
those of the turboprop and turbojet, assuming the use of the same gas gen- 
erator components in all three types of engines. 

The definitions of a turbofan (or ducted fan engine) and a bypass engine 
are quite confused, both in this country and abroad. For instance, the British 
say that a bypass engine is one in which the total airflow passes through the 
fan, after which the airflow divides into two portions (F igures 1-12 and I-13). 
This is in disagreement with many interpretations in the United States. The 
British also have their own definition for a turbofan or ducted fan. Even 
this is not used consistently. In the interest of being specific, Pratt & Whitney 
Aircraft has endeavored to standardize references to engines of these types. 

By Pratt & Whitney Aircraft defini- 
tion, a turbofan or a ducted fan en- 
gine is an engine having a duct-con- 
trolled, secondary airstream which 

was mechanically compressed by the 
Figure 1—12. Turbofan with fan as it entered the engine. Turbo- 

Mixed Exhaust fans are divided into engines with 
mixed exhaust streams (Figure 1-]2) 
and with nonmixed exhaust streams 
(Figure 1-13). Pratt & Whitney Air- 
craft defines a bypass engine as one 
in which the secondary airstream is 

Figure 1—13. Turbofan with Nonmixed subject to ram compression only 

(Figure 1-14). This arrangement re- 
quires that fuel be burned in the sec- 

ondary airstream in much the same 
manner as in an afterburner on a 
conventional engine to produce ad- 
ditional thrust. The performance 

Figure 1—14. Bypass Gas Turbine characteristics of a bypass engine 
Engine differ from those of the turbofan. 





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Gas Turbine Engine Components 

The physical features and functions of the major components which are in- 
cluded in the various types of gas turbine engines will be discussed in the order 
of their location, front to rear, on the engine. Later, in Section III, technical 
aspects relating to the aerodynamics, thermodynamics and performance of 
these same components will be discussed. 

Turboprop and Turboshaft Engine Nose Sections 

The configuration of a gas turbine engine dictates that whenever the engine is 
to drive a propeller or perform other work by means of a shaft, the drive shaft 
mechanism be placed at the front of the engine in order to leave the area at 
the rear clear for the engine jet exhaust. Because of the high operating rpm of 
a turbine engine, a reduction gear arrangement of considerable size is neces- 
sary to reduce the rotational speed of the turbine to usable propeller or shaft 
speed. The reduction is often on the order of about 9 to 1, which contrasts 
sharply with that of a reciprocating engine where a speed reduction of 2 or 3 
to 1 is typical. The large amount of horsepower which these engines are ca- 
pable of delivering and the requirement for a high reduction gear ratio pre- 
sent a real problem to engine design engineers. The reduction mechanism must 
be capable of handling heavy loads and yet must be both light in weight and 
small in frontal area. Adequate lubrication of the gearing must also receive 
special consideration. 

Figure 2-1 illustrates the nose 
section arrangement of the 
Pratt and Whitney Aircraft 
PT2 turboprop (or T34, as the 
military version of the engine is 
known). The nose section of 
this engine is more or less typi- 
cal of that of most others. In 

Figure 2—1. Engine Nose Section and 

SO SEs the nose section Propeller Reduction Gearing, P&2WA 
is offset from the turbine and Turboprop Engine (134) 

compressor shaft, leaving a 
clear entrance to the compressor air inlet. This is done, however, at some 
sacrifice to the minimum attainable frontal area of the engine. 





Air Inlet Duct 

Normally, the air inlet duct is considered an airframe part and not part of 
the engine. However, the duct, itself, is so important to engine performance 
that it must be considered in any discussion of the complete engine. 

A gas turbine engine consumes six to ten times as much air per hour as a recip- 
rocating engine of equivalent size. The air entrance passage is correspondingly 
larger. Furthermore, it is more critical than a reciprocating engine air scoop 
in determining engine and aircraft performance, especially at high airspeeds. 
Inefficiencies of the duct result in successively magnified losses through other 
components of the engine. The inlet duct has two engine functions and 
one aircraft function. First, it must be able to recover as much of the total 
pressure of the free airstream as possible and deliver this pressure to the face 
of the engine with a minimum loss or pressure differential. This is known as 
“ram recovery.” Secondly, the duct must uniformly deliver air to the compres- 
sor inlet with as little turbulence as possible. As far as the aircraft is concerned, 
the duct must hold to a minimum the drag effect, which it, itself, creates. 

Pressure drop or differential is caused by the friction of the air along both 
sides of the duct and by the bends in the duct system. Smooth flow depends 
upon keeping the amount of turbulence, as the air enters the duct, to a min- 
imum. The duct must have a sufficiently straight section to ensure smooth, 
even airflow within. The choice of configuration of the entrance to the duct 
is dictated by the location of the engine within the aircraft and the airspeed, 
altitude and attitude at which the aircraft is designed to operate. There are 
two basic types of inlet ducts, the single entrance duct and the divided 
entrance duct. 

Single Entrance Duct 

The single entrance type of duct (Figure 2-2) is the 
simplest and most effective because the duct inlet is 
located directly ahead of the engine and aircraft in 
such a position that it scoops undisturbed air. Also, 
the duct can be built either in a straight configuration 
or with only relatively gentle curvatures. In a single 
engine aircraft installation where the engine is 
mounted amidships, the duct is neces- 
sarily long. While some pressure drop 
is occasioned by the long duct, the con- 
dition is offset by smooth airflow char- 
acteristics. In multiengine installations, 
a short, straight duct, or one that is 
Figure 2—2. nearly straight, is a necessity. While 
this short, straight duct results in mini- 
mum pressure drop, the engine is apt to suffer from inlet turbulence, espe- 
cially at slow airspeeds and/or high angles of attack. 





Divided Entrance Duct 

The requirements of high-speed, single engine aircraft, in which the pilot 
sits low in the fuselage and close to the nose, render it difficult to employ the 
single entrance duct. Some form of a divided duct which takes air from 
either side of the fuselage may be required. This divided duct can be either 
a wing root inlet (Figure 2-3) or a scoop at each side of the fuselage (Figure 
2-4). Either type of duct presents more problems to the aircraft designer than 
a single entrance duct because of boundary layer problems and the difficulty 
of obtaining sufficient air scoop area without imposing prohibitive amounts 
of drag. Internally, the problem is the same as that 
encountered with the single entrance duct; that is, 
to construct a duct of reasonable length, yet with 
as few bends as possible. The wing root inlet on air- 
craft on which the wing is located fairly far aft, 
presents a design problem because, al- 
though short, the duct must have con- 
siderable curvature in order to deliver 
air properly to the compressor inlet. 
Scoops at the sides of the fuselage are 
often used. These side scoops are 
placed as far forward as possible in 
Figure 2—3. order to permit a gradual bend toward 
the compressor inlet, making the airflow characteristics approach those of 
a single entrance duct. A series of small rods is sometimes placed in the side 
scoop inlet to assist in straightening the incoming airflow and to pre- 
vent turbulence. 

With ducts of any type, careful 
construction is very essential. Good 
workmanship is also needed when 
an inlet duct is repaired. Surpris- 
ingly small amounts of airflow dis- 
tortion can result in appreciable 
loss in engine efficiency or can be 
responsible for otherwise unex- 
plainable engine compressor stall 
conditions. Protruding rivet heads 
or poor sheet metal work can play 
havoc with an otherwise accept- 
able duct installation. Figure 2—4. 


. JNSTR. 200 




Bellmouth Compressor Inlets 

Although not a duct in the true sense of the word, a bell- 
mouth inlet is usually installed on an engine being cali- 
brated in a ground test stand to lead the outside static 
air to the inlet guide vanes of the compressor. This type 
of inlet is easily attached and removed. It is designed 
with the single objective of obtaining very high aerody- 
namic efficiency. Essentially, the inlet is a bell-shaped Figure 2—5. 
funnel having carefully rounded shoulders which offer 

practically no air resistance (Figure 2-5). Duct loss is so slight that it is con- 
sidered zero. The engine can therefore be operated without the complications 
resulting from losses common to an installed aircraft duct. Engine perform- 
ance data such as rated thrust and thrust specific fuel consumption are ob- 
tained while using a bellmouth compressor inlet. Usually, the inlets are fitted 
with protective screening. In this case, the efficiency which is lost as the air 
passes through the screen must be taken into account when very accurate 
engine data are necessary. 

Turboprop Compressor Inlets 

The air inlet on a turboprop is more of a problem than that on a turbojet 
because the propeller drive shaft, the hub and the spinner must be considered 
in addition to the usual other inlet design factors. The ducted spinner arrange- 
ment (Figure 2-6) is generally considered the best inlet design of the turbo- 
prop engine as far as airflow and aerodynamic characteristics are concerned. 
However, the ducted spinner is heavier and is more difficult to maintain and 
to anti-ice than the more conventional streamline spinner arrangement which 
is more frequently used. A conical spinner, which is a modified version of the 
streamline spinner, is sometimes employed. In either event, the arrangement of 
the spinner and the inlet duct is similar to that shown in Figure 2-7. When the 
nose section of the turboprop engine is offset from the main axis of the engine, 

an underscoop arrangement similar to that in Figure 2-8 may be employed. 

Figure 2—6. Figure 2-7. Figure 2—8. 

Compressor Inlet Screens 

In connection with air inlet ducts, mention must be made of compressor inlet 
screens. The appetite of a gas turbine for nuts, small bolts, rags, small hand 
tools and the like is well known. To prevent the engine from readily ingesting 
such items, a screen is sometimes placed across the engine air inlet at some 
location along the inlet duct. The advantages and disadvantages of a 
screen of this type vary. If the engine is readily subject to internal damage, 


j eo as would be the case, for instance, of an engine having an axial compressor 
fitted with aluminum compressor blades, an inlet screen is practically a 
P “must.” Screens, however, add appreciably to inlet duct pressure loss and 

are very susceptible to icing. Failures due to fatigue are also a problem. A 
failed screen can sometimes cause more damage than no screen at all. In 
some instances, inlet screens are made retractable and may be withdrawn 
from the airstream after take-off or whenever icing conditions prevail. Such 
screens are subject to mechanical failure and add both weight and bulk to 
the installation. In large engines having steel or titanium compressors with 
blades which do not damage easily, the disadvantages of compressor screens 
outweigh the advantages, so they are not generally used. 



Tian ————— The combustion of fuel and air 
Wim ee at normal barometric pressure will 
not produce sufficient energy to en- 
able enough power to be extracted 
from the expanding gases to pro- 
duce useful work at reasonable ef- 


P ficiencies. The energy released by combustion is proportional to the mass of 
| air consumed. Therefore, more air is needed to increase the efficiency of the 

o combustion cycle than normal barometric pressure will provide. In both 
P reciprocating and gas turbine engines, the fuel/air mixture or air must be 
| compressed in order that a maximum amount of air can be handled in a given 
- volume. Compression in a reciprocating engine is accomplished by means 

of a piston which acts as a plunger, compressing the fuel/air mixture as the 

piston moves into a closely fitting tube which is closed at one end. The tube 
P is the engine cylinder. Many reciprocating engines further increase the 
| amount of air consumed by the addition of a supercharger. 

r The gas turbine must rely upon some other means of compression. Finding a 
| satisfactory manner in which to accomplish this necessary phase of the gas 
turbine cycle constituted the main stumbling block during the early years of 
turbojet engine development. Great Britain’s Sir Frank Whittle solved the 
problem by using a compressor of the centrifugal type. This form of compres- 
sor is still being used successfully in many of the smaller gas turbine engines 
today. However, the efficiency levels of single-stage centrifugal compressors 
are relatively low. The efficiencies of multistage centrifugal compressors are 
- somewhat better, but still do not compare with those of axial-flow compres- 
sors. A compression ratio of 4 or 5 to 1 is about the maximum capability of 
single-stage centrifugal compressors. Axial compressors, on the other hand, 
p produce much higher ratios. A high-efficiency dual axial compressor, for in- 
| stance, can attain a ratio of 12 to 1, or better. Axial compressors have the 
added advantages of being more compact and presenting a relatively small 
OT frontal area, which are important features in an aircraft engine. Therefore, 
most large gas turbines employ this type of compressor. 

= ae 




Centrifugal Compressors 

Centrifugal compressors operate by taking in outside air near their hub and 
rotating it by means of an impeller. The impeller, which is usually an alu- 
minum alloy forging, guides the air toward the outer circumference of the com- 
pressor, building up the velocity of the air by means of the high rotational 
speed of the impeller. The compressor consists of three main parts: an im- 
peller, a diffuser and a compressor manifold 
(Figure 2-9). Air leaves the impeller at high 
speed and flows through the diffuser which 
converts high-velocity kinetic energy to a 
low-velocity, high-pressure energy. The dif- 
fuser also serves to straighten the airflow 
and to turn. the air so that it may be picked 
up by the compressor manifold which acts 
as a collector ring. The diffuser blades direct Figure 2-9. 

the flow of air into the manifold at an angle 

designed to retain a maximum of the energy imparted by the impeller. They 
also deliver air to the manifold at a velocity and pressure which will be 
satisfactory for use in the burner section of the engine. 

The compressor shown in Figure 2-9 is known as a single- 
face or single-entry compressor. A variation of this is the 
double-face or double-entry compressor in which the im- 
peller is constructed as shown in Figure 2-10. The double- 
face compressor can handle the same amount of airflow 
and has a smaller diameter than a single-face compressor. 
This advantage is partially offset by the complications in- 
volved in delivering air from the engine inlet duct to the 
rear face of the impeller. Double-entry centrifugal com- 
pressors must have a plenum chamber to enable the in- 
coming air to be collected and fed to the rear impeller. 
Plenum chambers are, in essence, air chambers in which the 
compressor inlet air is brought to low velocity after having passed through 
the inlet duct of the aircraft. This air is brought in at ambient pressure plus 
ram pressure. The pressure in the plenum chamber is therefore greater than 
that of the outside atmosphere. The plenum chamber acts as a tool by means 
of which the rear impeller is able to receive its air supply. 

Figure 2—10. 

Multistage centrifugal compressors con- 
sist of two or more single compressors 
mounted in tandem on the same shaft 
(Figure 2-11). The air compressed by the 
first stage is passed on to the second 
stage at its point of entry near the hub. 
This stage further compresses the air be- 
fore passing it on to still another stage, 
if there is one. Figure 2—11. 



Axial Compressors 

rotating “rotor” blades and stationary “stator” vanes which are concentric 
with the axis of rotation. Unlike a turbine which also employs rotor blades 
and stator vanes, the flow path of an axial compressor decreases in Cross-sec- 



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af) o , 
ye Wi S4¢4Yy 

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f The air in an axial compressor flows in an axial direction through a series of 

) eV] 




o\Yed iE 
eink WA 

We NX 
yo : 
| Figure 2—12. Components and Assembly of Axial-flow Compressor 
P tional area in the direction of flow in proportion to the reduced volume of the 
| air as compression progresses from stage to stage. Figure 2-12 illustrates the 
rotor, the stator and the complete as- fe Tee 
r sembly of an axial compressor. Figure : 
2-13 shows a cross section of a typical Entrance 
axial compressor. an oS 
P —— Ti 
eo After being delivered to the compres- 
' sor by the air inlet duct, the incoming Figure 2—13. Axial-flow 
air passes through a set of inlet guide Compressor Cross Section 
vanes which prepare the flow for the 
F first-stage compressor rotor. Upon entering the set of rotating blades, the air, 

which is flowing in a general axial direction, is deflected in the direction of ro- 
tation. This change in the direction of flow is accompanied by a decrease in 
P velocity with a resultant rise in pressure through diffusion. The air is arrested 
and turned as it is passed on to a set of stator vanes, following which it is 
again picked up by another stage of rotating blades where the pressurizing 
P process continues. The stator vanes practically nullify the rotational effect 
given the air by the rotating blades. For example, air that enters the compres- 
sor at any particular position (for instance, 12 o'clock) continues on through 

P the entire engine, including the ‘et nozzle, without rotatin reatly, seldom 
J g greatly 

wandering off farther than 100° to 180°. It is the mechanics of the pressure 

- rise through diffusion, unassisted by centrifugal force, which constitutes the 


principal difference between axial-flow and centrifugal compressors. 

The airflow in an axial compressor is diffusing and, as a result, is very un- 

stable. High efficiencies can be maintained only at very small rates of dif- 

fusion. When compared with a turbine, quite a number of compressor stages 

p are necessary. Also, the permissible turning angles of the blades are consider- 
ably smaller than those which can be used in turbines. These are the reasons 
why an axial compressor must have many more stages than does the turbine 

im which drives it. In addition, more blades and consequently more stages are 
needed because the compressor, in contrast to a turbine, is endeavoring to 
push air in a direction that it does not want to go. 


. JNSTR. 200 




Aerodynamic principles are applied to compressor blade design in order to 
increase efficiency. The blades are treated as lifting surfaces like aircraft 
wings or propeller blades and are given an airfoil cross section, as shown in 
Figure 2-14. Unlike the wings and propeller blades, there is a cascade effect 
caused by one blade closely following an- 

DIRECTION other as the compressor rotates. Before 
the air can completely separate from the 
trailing edge of one blade, it is affected by 
the leading edge of the blade which fol- 
iia lows. The cascade effect is ‘a primary con- 
of blade sideration in determining the airfoil sec- 
eer tion, angle of attack, and the spacing be- 
Side View tween blades to be used for the compres- 
“a sor blade design for any given axial com- 

pressor. The blades must be designed to 

Figure 2—14. Airfoil Cross Section withstand the high centrifugal forces as 
of Axial Compressor Rotor Blade well as the aerodynamic loads to which 
they are subjected. Consideration must 

also be given the vibration caused by the impact of air at high velocity 


Apparent angle 
of attack 


Airfoil section 

_ Issuing in small streams from the spaces between blades. High efficiencies are 

attained only when tolerances within the compressor are held to a minimum. 
Of these, the clearance between the rotating blades and their outer case is very 
important. For this reason, some compressor blades are made with a knife- 
edge tip. The assembled compressor rotor fits easily into the compressor case. 
However, as the blades expand from the heat generated within the com- 
pressor, they grow longer and rub against the compressor case. As they do so, 
they grind off at the knife-edge tips and establish their own clearances. It is 
thus assured that each blade will fit the case as snugly as possible. 

The compressor rotor drum of an axial flow compressor may be machined 
from a solid forging or built up from individual discs. Rotating compressor 
blades are cast or forged from aluminum, steel or titanium and are fastened 
in slots around the periphery of the discs. The stationary vanes are attached 
to the compressor casing between each adjacent rotor stage. The rotor assem- 
bly turns at extremely high speed and must be rigid, well aligned and well 
balanced. Each rotor stage is usually statically balanced at the time of assem- 
bly. The completed rotor unit is then placed on a balancing machine and is 
accurately checked for dynamic balance. 

Dual Axial Compressors 

A single axial compressor might theoretically be built to consist of as many 
stages as would be necessary to produce any required compression ratio. If 
such were the case, at some specific compressor speeds the rearmost stages of 
the compressor would operate inefficiently and the foremost stages would be 
overloaded. Such a condition would produce compressor stall. The condition 
could probably be corrected, in part, by bleeding interstage compressor air 
overboard during part-throttle operation. Excessive air bleeding, however, is 
wasteful. Greater flexibility for part-throttle conditions and for starting can be 
attained more efficiently by splitting the compressor into two mechanically 


a a =o 




independent rotor systems. Each 
is driven by its own separate tur- 
bine and each assembly is free to 
rotate at its own best speed (Fig- 
ure 2-15). The high pressure com- 
pressor has shorter blades than 
the low pressure compressor, and 
is lighter in weight. Since the 
work of compression by the high Figure 2—15. Dual Axial Compressor 
pressure compressor heats the air System 

within the compressor to higher 

temperatures than occur within the low pressure compressor, higher. tip 
speeds are possible before the blade tips attain their limiting Mach number 
because of the fact that the speed of sound increases as the air temperature 
increases. Hence, the high pressure compressor can run at a higher speed 
than the low pressure compressor. 

Low Pressure Compressor and Turbines 

High Pressure 
Compressor and Turbine 

f fa 

When dual, or, as sometimes called, “split,” compressors are used, high com- 
pression ratios can be attained with minimum total compressor weight and 
frontal area. Usually, the rear, or high pressure, compressor rotor is “speed- 
governed” by the engine fuel control and is the rotor to which the engine 
starter drive is connected. Only part of the complete compressor, and the 
lighter part, at that, is cranked, which considerably reduces the torque re- 
quired to start the engine. The size and weight of the starting system may 
therefore be appreciably less. With the rear, high pressure compressor turning 
at governed speed, the front, or low pressure, compressor is rotated by its 
turbine at whatever speed will ensure optimum flow through the compressor. 
The components of the compressor adjust themselves to part-throttle opera- 
tion, with a minimum of interstage bleeding, to prevent stall or surge. Flow 
matching is assured between the compressors and the turbines throughout the 
operating range of the engine. With the front and rear rotors working in 
harmony instead of interfering with each other, compression ratio can be in- 
creased without decreasing efficiency. 

The nature of a dual axial compressor leads to still further advantages. With 
the high pressure compressor governed to constant speed, there is a change in 
speed of the low pressure compressor as changes occur in the temperature of 
the compressor inlet air. The speed increases as the temperature becomes 
colder. This is due to the fact that the power required to compress cold air to 
a given pressure or pressure ratio is less than the power required to compress 
warm or hot air. With less work to do, the turbine for the low pressure com- 
pressor turns faster. The low pressure compressor speeds up correspondingly. 
The resulting increased airflow has an effect upon the high pressure turbine 
and compressor. However, high pressure compressor speed is controlled 
through the fuel control. This speed control, in turn, tends to limit the energy 
delivered to the low pressure turbine, so that equilibrium is finally established, 
the high pressure compressor operating at controlled speed and the low pres- 
sure compressor operating at a speed that, on a cold day or at altitudes with 





cold temperatures, is somewhat higher than at standard day sea level tem- 
peratures. On a hot day, the low pressure compressor will operate at a slower 
speed than on a standard day. 

When constant turbine speed is maintained with varying compressor inlet 
temperature, the gas temperature entering the turbine increases as the com- 
pressor inlet temperature gets colder than standard day sea level temperature 
and decreases as the compressor inlet temperature gets warmer. This trend 
causes the metal parts to run hotter on cold days. To obtain more favorable 
operating conditions for the hot parts of the engine and also more thrust, the 
speed control in the fuel control of most dual axial compressor engines is 
biased for compressor inlet temperature so that compressor speed is increased 
for warm inlet temperatures. Although this applies only to the high pressure 
compressor, it also has a tendency to stabilize the speed of the low pressure 
compressor. Thus, as altitude is gained and the low pressure compressor 
speed increases with colder inlet air temperatures, the loss of thrust with alti- 
tude, due to reduced air density, is less than one might expect. This variation 
in compressor speed with changing inlet temperature, which is known as 
“speed bias,” is explained in greater detail later. 

Most single axial compressors are designed with the outside diameter of all 
of the rotor stages held either constant or nearly constant, reduced airflow 
area being obtained by increasing the inside diameter. This results in a con- 
stant over-all engine diameter. High rotor blade tip speeds and short blades 
at the high pressure end of the compressor are possible because the outside 
diameter at the compressor exit is large. With the dual type of axial com- 
pressor, the diameter of the high pressure rotor can be considerably reduced 
over that of a correspondingly large, single axial compressor. Because it is 
possible for the high pressure compressor to have a smaller diameter than 
that of the low pressure compressor, the blade tip speeds may be more easily 
held at their limiting Mach number at high rotor speeds while still being light 
in weight. If the compressor rotors decreased in external diameter (the inside 
diameter being held constant), the blades at any given stage would have to 
have a greater length in order to sweep the same area than could be the case 
with constant diameter rotors. The constant diameter design permits more 

efficient blade shape. With the rel- 

atively large outside diameter and 
the shorter blades, the blade roots 
are placed farther out (at a greater 

diameter) on the rotor. This design 

permits that portion of the high pres- 

sure compressor blade near and at 
the root to do more of its share of the 
work than is possible in a compres- 
sor which has long blades. By hav- 

ing a small and constant external Figure 2—16. Cross-sectional Area 
diameter, the shape and size of the Comparison of Single and Dual 
high pressure rotor permits a “wasp- Axial Compressor Engines 


| r PWA OPER. [NSTR. 200 

en waist” engine configuration. This is highly desirable because it enables the 
fuel control and other outside engine parts and accessories, such as the 
starter, to be mounted in the high pressure compressor area without adding 
appreciably to the over-all engine frontal area (Figure 2-16). 

Compressor Stall and Compressor Airbleed 

f It is a characteristic common to gas turbine compressors of all types, including 
the compressor of a reciprocating engine turbosupercharger, to stall under 
f certain operating conditions. Some call this surge. Others endeavor to differ- 
entiate between stall and surge, claiming that the two terms may be considered 
distinct. For all practical purposes, however, the two terms may be considered 
f analogous and may be treated as one and the same thing. The difference in 
symptoms, if any, is so obscure that detection is either difficult or impossible. 
The stall characteristic is especially noticeable in high compression ratio axial 
r compressors, both single and dual rotor. Compressor stall occurs in many 
different forms and under many different conditions. It is neither easy to 
r describe nor to understand, particularly because the stall characteristics of 
no two engine designs will be the same. Suffice it to say here that, in general, 
r stall results from an unstable air condition within the compressor. 

This unstable condition is often caused, in part, from air piling up in the rear 
stages of the compressor. When a compressor is not operating at its optimum 
rpm, especially at high altitude, the forward compressor blades may not be 
able to bite off enough air to be able to compress it sufficiently to force it on 

eo through the rear stages of the compressor. Insufficient pressure ratio across 
the compressor results, and the rear stages tend to become choked. The com- 
pressor blades, which are, themselves, miniature airfoils, stall in much the 
same manner as does an aircraft wing at too slow a flying speed. 

In its milder form, compressor stall can be recognized by the condition 
known as “choo-choo,” which is occasionally encountered during ground 
engine operation at low thrust. In flight, under severe conditions of “jam” 
acceleration or when slipping or skidding during evasive action or when fly- 
ing in very turbulent air (which causes upset air conditions at the aircraft 
engine air inlet), stall may become sufficiently pronounced to cause loud 
bangs and engine vibration. In most cases, this condition is of short duration 
and will either correct itself or can be corrected by retarding the throttle to 
Idle and advancing it again, slowly. This subject will be discussed in more 
detail later. 

schedule at the expense of efficiency. The loss, however, would be so great 
that it is usually considered better to design the engine and the fuel flow sched- 
ule of the fuel control in such a manner that stall will normally be avoided 
during periods of acceleration, deceleration and low thrust operation. Among 
other things, to minimize the tendency of a compressor to stall, the com- 
pressor must be frequently “unloaded” during certain operating conditions by 
reducing the pressure ratio across the compressor for any given airflow. One 
ao method of doing this is by bleeding air from the middle or toward the rear of 
the compressor. In dual axial compressor engines, air is often bled from be- 


‘ Stall can be completely “designed out” of an engine fuel control operating 



tween the low and the high pressure compressor. Airbleed ports are located 
in the compressor section. These ports are fitted with automatic overboard 
bleed valves which usually are controlled by engine rpm. Frequently, the rpm 
at which the bleed valves open or close is varied automatically as a function 
of compressor inlet temperature. The bleed valves serve to facilitate engine 
starting and prevent compressor stall at high altitudes by ducting compressor 
air overboard during low thrust operation. The bleeds automatically close 
during high thrust operation. When the bleeds are open, they increase air- 
flow through the upstream portion of the compressor and reduce the pres- 
sure on the bottleneck downstream. These overboard airbleeds should not 
be confused with airbleeds which serve to provide low or high pressure 
air for engine service functions and to operate auxiliary equipment in the air- 
craft. These service bleeds also have effects upon compressor stall, which 
may occasionally be noted when airbleed-operated, aircraft or engine equip- 
ment is in use. Improved design, especially in dual axial compressors, is be- 
ginning to reduce the need for overboard airbleed and, in some engines, will 
eliminate this requirement altogether. 


The air leaving a compressor passes 
through a diffuser. The manner in 
which this is accomplished in a cen- 
trifugal compressor has already been 
described. The diffuser prepares the 
air for entry into the burners at low velocity by converting velocity (kinetic) 
energy into pressure energy. In both single and dual axial compressor en- 
gines, diffusion is achieved by compressor exit guide vanes located imme- 
diately aft of the rearmost compressor stage in what is known as the diffuser 
section or case of the engine, just forward of the combustion chamber. 

Fuel Manifolds and Nozzles 

Fuel is introduced into the air- 
stream at the front of the burners 
in spray form, suitable for rapid 
mixing with air and combustion. 
The fuel is carried from outside 
the engine, by a manifold system, to nozzles mounted in the burner cans. 
The most common type of nozzle employs a pressure atomizing system which 
ensures a finely atomized spray of uniform distribution throughout the range 
of fuel flows encountered during engine operation. Swirl-type nozzles are 
usually used to provide high flame speed with low axial air velocity. Multiple 
nozzles are frequently employed to handle large quantities of fuel with even 
distribution and to minimize the variation which might be caused by a clog- 
ging of any one of the nozzles. A primary and secondary fuel manifold and 
nozzle system is often used on large engines. These are sometimes called the 
pilot and main manifolds. The primary, or pilot, system provides sufficient 

oo Se 


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niet aa 


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PWA OPER. | NSTR. 200 

fuel for low thrust operation. At high thrust operation, the secondary, or 
main, system cuts in, and fuel commences to flow from both the primary 
and secondary elements of a double orifice nozzle. In a nozzle of this type, 
primary fuel is sprayed through a single orifice in the center of the nozzle. 
Secondary fuel is sprayed through a number of orifices spaced in a ring 
around the orifice of the primary manifold. 

Burner Section 

nae The burner section which contains 
Billi the combustion chamber is designed 
will to burn a mixture of fuel and air and 
; to deliver the resulting gases to the 
turbine at a temperature which will 
not exceed the allowable limit at the turbine inlet. The burners, within a very 
limited space, must add sufficient heat energy to the gases passing through 
the engine to accelerate their mass enough to produce the desired thrust for 
the engine and power for the turbine. The heat released per cubic foot of 
combustion space in a large turbojet engine is several thousand times as 
great as the heat released per cubic foot of burner space in an ordinary, 
home-heating oil burner. It is also interesting to note that the pressure within 
a 10,000 pound thrust turbojet combustion chamber, which is enclosed by 
only a relatively small thickness of steel wall, is approximately ten times 
as great as the pressure within the average industrial furnace which is en- 
closed by very thick walls of firebrick and other materials. 


The criteria for an acceptable burner are that the pressure loss as the gases 
pass through the burner must be held to a minimum; the combustion efficiency 
must be maintained at a high level; and there must be no tendency for the 
burner to blow out. No burning should occur after the gases leave the burner 
outlet, which means that complete combustion must take place entirely within 
the burner. The gases must have satisfactory temperature distribution and ac- 
ceptable maximum temperature as they enter the turbine. Fortunately, it has 
been found that a burner design which will satisfy these conditions and will 
be suitable for continuous operation will also have satisfactory engine start- 
ing characteristics. 

Combustion Chambers 

Combustion chambers may be either 
of the can, the annular or the can- 
annular type. For all of these, the 
design is such that less than a third 
of the total volume of air entering 
the chamber is permitted to mix with the fuel. The ratio of total air to fuel 
varies among different types of engines from 40 to 80 parts of air to one 
of fuel, by weight. A ratio of 50:1 is about average. Of the 50 parts of air, 
however, only about 15 parts, by weight, are used for burning. Therefore, 
all of the air in excess of these 15 parts bypasses the fuel nozzles and is used 
downstream to cool the burner surfaces and to mix with and cool the burned 
gases before they enter the turbines. 



a — ~ 



Can-type Burner and Multiple Combustion Chamber 

A can-type of burner arrangement is most 
frequently employed on centrifugal com- 
pressor engines. In this system, the air is 
divided as it leaves the diffuser and is 
ducted to individual combustion cans, or 
cylinders, arranged around the circum- 
ference of the burner section of the engine 
(Figure 2-17). Each burner or can con- 
tains its own fuel nozzle and burner liner 
(Figure 2-18). Primary air, introduced at 

the nozzle, serves to support the initial Figure 2—17. Typical, 
phase of combustion. Cooling, or second- Multiple-can-type Combustion 
ary, air passes between the liner and the Chamber Assembly 

burner case. The liner is provided with 
several series of holes or axial slots through which some of the secondary 
air enters from the outside. These holes furnish a layer of cooling air, flow- 
ing along the inside of the liner, and supply additional air for the combus- 
tion process. After combustion is 
completed, the heated airstreams 
from the multiple burners converge 
immediately upstream of the tur- 
, bine. At this point, the secondary 
joa air which bypassed the fuel nozzles 
mixes with the products of combus- 
Figure 2—18. Individual Can-type tion, cooling them to a temperature 

Burner which the turbine can tolerate. 



_ The can-type of burner has many advantages. Because of the smaller diameter 

of each unit as compared with that of the annular type of combustion cham- 
ber, structural strength is afforded, combined with light weight. This burner 
has excellent serviceability. Individual units may be removed from the engine 
for inspection or replacement without disturbing the engine installation. On 
the other hand, the design is such that the can diameter must be held to a min- 
imum, which means that the can type of burner does not lend itself to a short 
length. Furthermore, the fuel nozzle must be located in the center of the burn- 
er, which compels the air to travel a considerable distance into the can before 
it reaches the fuel, mixes with it and burns. Also, if trouble develops with any 
one nozzle, the turbine nozzle guide vanes are subjected to severe temperature 
differentials which can cause vane distortion. 

Annular Combustion Chamber 

Some axial compressor engines have a single annular combustion chamber 
similar to that shown in Figures 2-19 and 2-20. The liner of this type of burn- 


. PWA OPER. [NSTR. 200 

er consists of continuous, circular 
inner and outer shrouds around 
the outside of the compressor 
drive shaft housing. The liner is 
sometimes referred to as a “burner 

a basket” because it is perforated 
with holes which give it a basket- 

like appearance. Holes in the 
P shrouds allow secondary cooling 

air to enter the center of the com- 
bustion chamber. In the annular 
combustion chamber, fuel is in- 
troduced through a series of noz- 
zles at the upstream end of the 

- Figure 2—19. Typical Annular Combustion |. a 
| Citnilies liner. Because of their proximity 
to the flames, all types of burner 
‘a liners are short-lived in comparison with other engine components, requiring 
| frequent inspection and, sometimes, replacement. 
This type of burner has the advantage of being able to use the limited space 
r available most effectively, permitting better mixing of the fuel and air within a 
relatively simple structure. 
- An optimum ratio of burner Fuel Manifold Sekar Wana tial 
inner surface area to vol- Screen Inner Burner Shroud 
eo ume is provided, thus en- 
Pr suring maximum cooling of 

the gases as combustion takes 

place. The design also tends 
fF to prevent heat warping. 
However, the burner liner 
usually cannot be disassem- 

Primary Air Inlets Secondary Air Inlet 
Primary Compressed Air 

: : ——ammmnt Secondary Compressed Air 
| bled without removing the en- 

gine from the aircraft, which Figure 2—20. Cross Section of Annular 
- is a distinct disadvantage. Combustion Chamber 


Can-annular Combustion Chamber 

P In the can-annular arrangement, individual 
cans or liners are placed side by side in an 
annular chamber (Figure 2-2]). The cans 
P are essentially individual annular burners 
(Figure 2-22). Each is provided with an in- 
dividual concentric tube which substantially 
increases the effective burner length with- 
out adding to the physical dimensions. A 
e cluster of several nozzles is placed around 
| the perimeter of the forward end of the 

can. The structure is relatively small in 


rm diameter and has an inherent resistance to 
buckling when subjected to high pressure Figure 2—21. 



at operating temperatures. Special baffling is employed to swirl the combus- 
tion airflow and to give it a reverse direction. Satisfactory cooling is made 
possible by introducing boundary layer air through annular apertures located 
at suitable interyals along the —— of the can. The can-annular combustion 
| = = 7 chamber combines the advan- 
tages of both the can and the 
annular types and eliminates 
many of their disadvantages. 
A removable or telescoping 
shroud covers the entire burn- 
er section and permits reason- . 
ably easy access for inspection J 
Figure 2—22. or replacement of the cans 
without removal of the engine 
from the aircraft. The short burner length possible with a can-annular com- j 
bustion chamber prevents an excessive drop in the pressure of the gases 
between the compressor outlet and the flame area. The design provides an 
even temperature distribution at the turbine inlet without the danger of hot J 
spots occurring should one of the nozzles clog. 

ee Ga 



EP me 

The turbine extracts kinetic energy | J 
from the expanding gases which J 
FA] ) flow from the combustion cham- | 
‘: : ber, converting it into shaft horse- = 

power to drive the compressor and 
the accessories. In a shaft or propeller turbine, the turbine also supplies addi- 
tional power to perform whatever work may be required of the engine. 7 
Nearly three fourths of all of the energy available from the products of com- 
bustion is necessary to drive the compressor. If the engine is a turboprop or . 
turboshaft, the turbine is designed to extract all of the energy possible from 
the gases passing through the engine. So efficient is the turbine, in this case, 
that in a turboprop aircraft the propeller provides approximately 90 per cent d 
of the propulsive force, leaving but 10 per cent to be supplied by jet thrust. 

The axial-flow turbine is comprised of two main elements: a turbine wheel or 

rotor and a set of —— vanes (F were 2-23). The stationary section con- = 
7 .% 

" My, "NN D x ) J 

+" 4, | ) 4, 

Figure 2—23. Turbine Elements d 
sists of a plane of contoured vanes, concentric with the axis of the turbine and | 
set at an angle to form a series of small nozzles which discharge the turbine we 



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PWA OPER. | NSTR. 200 

gases onto the blades of the turbine wheel. For this reason, the stationary vane 
assembly is usually referred to as the turbine nozzle, and the vanes, them- 
selves, are called nozzle guide vanes. Turbine nozzle area is a critical part of 
turbine design. If too large, the turbine will not operate at its best efficiency. If 
too small, the nozzle will have a tendency to choke under maximum thrust 
conditions. The jets of escaping gases which are formed by the nozzle dis- 
charge are directed against the rotating turbine blades in a direction which en- 
ables the kinetic energy of the gases to be transformed into mechanical 
energy which is generated by the rotating turbine wheel. 

. <6 . . i i = t 
Turbines are divided into three types: impulse, kre acing Picardy orn 
reaction, and a combination of these two known 
as reaction-impulse. In the impulse type of tur- e 
bine, there is no change in pressure between the Turbine YY 
; : : ‘otation 

rotor gas inlet and the rotor gas exit (Figure 
2-24). The nozzle guide vanes are shaped to Rotor Gas Exit 

; : z Impulse Turbine 
form passages which increase the velocity and Rotor Blades 

reduce the pressure of escaping gases. In the re- Figure 2-24. Impulse 
action type of turbine, the nozzle guide vanes do Turbine Rotor Blades 
little else than alter the direction of gas flow to Tangential Velocity 
that required by the turbine wheel. The reduction eee 
in pressure and the increase in velocity of the 
gases are accomplished by the shape of the pas- yn 
sage between the rotor blades (Figure 2-25). In a Turbine 
‘ ‘ ; ‘ Rotation 
gas turbine engine, the turbine is usually a bal- 

anced combination of both of these two types and SSN 

is known as a reaction-impulse turbine. The design agtecns Apt avenh 
is intended to achieve both a desirably small di- Figure 2—25. Reaction 
ameter and a proper match with the compressor. Turbine Rotor Blades 

Turbines may be either single- or multiple-stage. When the turbine has more 
than one stage, stationary vanes are inserted between each rotor wheel and 
the rotor wheel downstream, as well as at the entrance and exit of the turbine 
unit. Each set of stationary vanes forms a nozzle vane assembly for the turbine 
wheel which follows. The exit set of vanes serves to straighten the gas flow 
before passage through the jet nozzle. The wheels may or may not operate 
independently of one another, depending upon the type of engine and the 
power requirements of the turbine. 

Shaft rpm, gas flow rate, turbine inlet and outlet temperature and pressure, tur- 
bine exhaust velocity, and the required power output must all be given con- 
sideration by the designer of the turbine. If the engine is equipped with a dual 
compressor, the problem is more complex than ever, since the turbine also 
must be dual or “split.” In this event, the forward part of the turbine, which 
drives the high pressure compressor, can be single-stage because it receives 
gases of high energy directly from the burner and turns at a higher rpm than 





the turbine for the low pressure compres- High Pressure J 
sor. By the time that the gases reach the —— 
back part of the turbine, which drives 

the low pressure compressor, they have Low Pressure 
expanded in passing through the first tur- oad 
bine, and considerably more blade area is 

needed if proper work or energy balance 

is to be maintained. To accomplish this, 

a two-stage turbine is employed for the Figure 2—26. Two-part Turbine 
second part of the turbine (Figure 2-26 ‘3 for Dual Compressor Engine 

To be capable of supplying sufficient power for the compressor, the turbine 
must be designed so that the gases have a high expansion ratio. This, in turn, 
results in a large temperature drop of the gases passing through the turbine 
and a cool turbine exhaust. If the engine is equipped with an afterburner, a 
cool exhaust enables more fuel to be burned in the afterburner without ex- 
ceeding the temperature limit of the construction materials used in the after- 
burner. As a result, the afterburner thrust can be increased considerably above 
the maximum thrust allowed if the exhaust gases were very hot. 

The turbine wheel is a dynamically 

balanced unit consisting of steel al- 

Shroud loy blades, or buckets, as they are 

sometimes called, attached to a ro- 

tating disc. The base of the blade is 

Fir Tree Base usually of a so-called “fir tree” de- 

sign to enable it to be firmly attached 

to the disc. In some turbines, the 

rotating blades are open at their 

outer perimeter. In others, the blade 

is shrouded at the tip, as shown in 

Figure 2—27. Shrouded Turbine Figure 2-27. The shrouded blades 

Rotor Blades form a band around the perimeter 

of the turbine wheel which serves to 

reduce blade vibrations. The weight of the shrouded tips is offset because the 

shrouds permit thinner, more efficient blade sections than are otherwise pos- 

sible because of vibration limitations. Also, by acting in the same manner 

as aircraft wing tip fences, the shrouds improve the airflow characteristics 
and increase the efficiency of the turbine. 

Turbines are subjected to both high speeds and high temperatures. High 
speeds result in high centrifugal forces and because of high temperatures tur- 
bines must operate close to temperature limits which, if exceeded, will lower 
the strength of the construction materials used in turbines. Turbine blades 
and vanes tend to change pitch slightly with continued use, straightening to- 
ward low pitch. They also undergo distortion or lengthening of the blade or 
vane which is known as “creep.” Creep means that the blade stretches or 
elongates. This condition is cumulative, the rate of creep being determined 


— | 


—_ [67 




by the load imposed on the turbine and the strength of the blade, which is 
determined by the temperature within the turbine. Since changes in pitch and 
creep are more pronounced if engine operating limits are not treated with re- 
spect, it behooves the pilot or flight engineer who strives for long and depend- 
able engine life, to observe closely the temperature and rpm limits stipulated 
by the engine manufacturer and published in the aircraft Flight Handbook. 
The subject of creep as it affects engine life is discussed in more detail later. 

Exhaust Ducts 

| The term, “exhaust duct,” applies to 
el ill z the engine exhaust pipe or tailpipe 
connecting the turbine outlet and the 
jet nozzle of a nonafterburning en- 
gine. Although an afterburner might 
also be considered a type of exhaust duct, afterburning is a subject in itself 
and is dealt with subsequently. 

If the engine exhaust gases could be discharged directly to the outside air in 
an exact axial direction at ‘the turbine exit, an exhaust duct might not be 
necessary. This, however, is not practical. A larger total thrust can be ob- 
tained from the engine if the gases are discharged from the aircraft at a higher 
velocity than is permissible at the turbine outlet. An exhaust duct is therefore 
added both to collect and straighten the gas flow as it comes from the turbine 
and to increase the velocity of the gases before they are discharged from the 
exhaust nozzle at the rear of the duct. Increasing the velocity of the gases in- 
creases their momentum and increases the thrust produced. 

An engine exhaust duct is often referred to as the engine tailpipe. Although 
the duct, itself, is essentially a simple stainless steel conical or cylindrical 
pipe, the assembly to the rear of the turbine usually also includes an engine 
tail cone and struts inside the duct. The tail cone and the struts serve to add 
strength to the duct, to impart an axial direction to the gas flow and to smooth 
the flow. Immediately aft of the turbine outlet and usually just forward of the 
flange to which the exhaust duct is attached, the engine is instrumented for 
turbine discharge pressure. In large engines, it is not practical to measure the 
internal temperature at the turbine inlet, so the engine is often also instru- 
mented for exhaust gas temperature at the turbine outlet. One or more pres- 
sure probes are inserted in the exhaust duct to provide adequate sampling of 
the exhaust gases. Thermocouples are also inserted in the duct to measure the 
temperature of the gases coming from the turbine. The gradually diminishing 
cross-sectional area of a conventional convergent type of exhaust duct is 
capable of keeping the mass flow through the duct constant at velocities not 
exceeding Mach 1 at the exhaust nozzle. 






Conventional Convergent Exhaust Nozzle 

The rear opening of the exhaust duct 
(Figure 2-28) is the exhaust nozzle. The 
nozzle acts as an orifice, the size of which 
determines the density and velocity of the 
gases as they emerge from the engine. In 
most nonafterburning engines, this area is 
quite critical and, for this reason, is fixed Figure 2—28. Conventional 

at the time of manufacture. The area Convergent Exhaust Duct 
should not be altered unless provision has 

been made to do so. Adjusting the area will change both the engine perform- 
ance and the exhaust gas temperature. Some engines, however, are trimmed 
to their correct rpm or exhaust gas temperature by means of altering the ex- 
haust nozzle area. When this is the case, small tabs which may be bent, as re- 
quired, are provided on the exhaust duct at the nozzle opening, or small, ad- 
justable pieces called “mice” are fastened, as needed, around the perimeter of 
the nozzle to change the area. Occasionally, engines are equipped with vari- 
able-area nozzles which are opened or closed, usually automatically, as fuel 
flow is increased or decreased. The velocity of the gases within a convergent 
exhaust duct is usually held to a subsonic speed. Although the velocities at 
the nozzle may approach Mach 1, losses in efficiency will be incurred if the 
velocity of the gases becomes sonic before they leave the engine tailpipe. 

Exhaust Nozzle 

Convergent-divergent Exhaust Nozzle 

Subsonic Supersonic 
Convergent Divergent 

Whenever the pressure ratio across an 
Section Section 

exhaust nozzle is high enough to pro- 
duce gas velocities which might ex- 
ceed Mach 1 at the engine exhaust 
nozzle, more thrust can be gained by 
using a convergent-divergent type of 
nozzle, provided the weight penalty 
is not so great that the benefit of the ; j 
additional thrust is nullified. The ad- *i9¥"e 2—29. Convergent-divergent 

; Exhaust Duct (Nozzle) 
vantage of a convergent-divergent 
nozzle (or simply a “C-D nozzle,” as it is more commonly called) is greatest 
at high Mach numbers because of the resulting higher pressure ratio across 
the engine nozzle. If the pressure ratio through a subsonic exhaust duct is 
great enough, as will be the case when the pressure at the entrance to the ex- 
haust duct becomes approximately twice that at the exhaust nozzle, the change 
in velocity through the duct will be sufficient to cause sonic velocity (Mach 
1) at the nozzle. At very high flight Mach numbers, the pressure ratio be- 
comes much more than 2.0 and if a C-D nozzle were used, the velocity at the 
exhaust nozzle would become correspondingly greater than Mach 1. This is 
a distinct advantage, provided the nozzle can effectively handle these high 
velocities. A conventional convergent exhaust duct and nozzle, however, can- 
not tolerate velocities in excess of Mach 1, and losses in efficiency occur at 
high pressure ratios. 

Gas Attains Sonic Velocity 


Exhaust Nozzle 

a _ & Ge 




At sonic and supersonic flow rates, the rate of change in volume of a gas is 
greater than the rate of change in velocity, whereas, at subsonic flow rates, the 
rate of change in volume is proportional to the rate of change in velocity. 
After sonic velocity is attained, the gas expands more rapidly than it acceler- 
ates. To ensure that a constant weight or volume of gas will flow past any 
given point after sonic velocity is reached, the rear part of the duct must be 
enlarged to accommodate this additional weight or volume. If this were not 
done; the nozzle would choke. This section of the duct is known as divergent. 
The rate of increase in area in a divergent duct is just sufficient to allow for 
the increase in the rate of change in volume of the gases after they become 
sonic. The rate of increase in area is not sufficient to handle the supersonic 
flow at a constant rate of flow. The pressure therefore continues to increase 
as the gases proceed down the duct, increasing in a manner similar to the 
pressure buildup through a subsonic conventional convergent duct. This 
results in an increase in velocity through the duct, which proportionally 
matches the increase in weight or volume of fhe gases at any given point along 
the duct. Thus, when the increase in area of the duct is accomplished effec- 
tively, the weight of the gas flow past any given point will remain constant, 
and steady, mass flow conditions will be maintained. 

When a divergent duct is employed in combination with a conventional 
exhaust duct, the type is called a convergent-divergent exhaust duct (Figure 
2-29). Since the entire configuration is, in effect, considered a nozzle, the duct 
is often referred to as a convergent-divergent exhaust nozzle. In the conver- 
gent-divergent (or “C-D”) nozzle, the convergent section is designed to handle 
the gases while they remain subsonic and to deliver the gases to the throat of 
the nozzle just as they attain sonic velocity. The divergent section handles the 
gases, further increasing their velocity after they emerge from the throat and 
become supersonic. 

In the discussion on thrust (Section I), it was pointed out that all of the pres- 
sure generated within an engine cannot be converted to velocity, particularly 
when a convergent nozzle is used. The additional pressure results in additional 
thrust which, as has been shown, must be added when the total thrust 
developed by the engine is computed. The additional thrust is developed 
inefficiently. It would be much better to convert all of the pressure within the 
engine to velocity and thus develop all of the engine thrust by means of 
changes in momentum. In theory, at least, a C-D nozzle does this and, because 
it develops this additional part of the total thrust more efficiently, it enables 
an engine to produce more total net thrust than the same basic engine would 
generate if it were equipped with a conventional convergent duct and nozzle. 
Production of additional engine thrust by the use of a C-D nozzle will become 
increasingly important as routine aircraft flight speed continues to progress 
into realms well above Mach 1.0. 

The C-D nozzle would be nearly ideal if it could always be operated under 
the exact conditions for which it was designed. However, if the rate of change 
in duct area is either too gradual or too rapid for the calculated increase in 
weight of the gases, unsteady flow downstream of the throat will occur with 
an accompanying loss of energy which ultimately means loss of thrust. If the 





rate of increase in area of the duct is too little, the maximum gas velocity 
which can be reached will be limited. If the rate of increase is too great, the 
gas flow will break away from the surface of the nozzle and the desired in- 
crease in velocity will not be obtained. As the exhaust gases accelerate or de- 
celerate with changing engine and flight conditions, their pressure fluctuates 
above or below the pressure ratio for which the nozzle was designed. When 
this occurs, the nozzle no longer converts all of the pressure to velocity, and 
the nozzle begins to lose much of its purpose. 

The solution to this dilemma is a C-D nozzle with a variable cross-sectional 
configuration, which can adjust itself to changing pressure conditions. 
Although a nozzle of this type tends to be too bulky, heavy, and complicated 
to be practical, progress is being steadily made in solving the problems in- 
volved. It is entirely possible that C-D nozzles with an automatically variable 
area may become quite commonplace on engines of the future. Meanwhile, 
fixed-area, C-D nozzles, designed for best performance at some predetermined 
altitude and airspeed, are now coming into use. These nozzles provide enough 
additional thrust at the altitude and ‘airspeed at which the aircraft is expected 
to perform its basic mission to justify their use at other altitudes and airspeeds. 
Such nozzles are equipped with two-position flaps which are open when the 
afterburner is operating and are closed when it is not. 

Although there is no reason why a C-D nozzle may not be used to advantage 
on a nonafterburning engine, their present use is confined to afterburning 
engines in fighter-type aircraft, these being the only aircraft currently operat- 
ing at sufficiently high Mach numbers to make the use of a C-D nozzle worth- 
while for the additional weight involved. Jet nozzle geometry has a significant 
effect on the thrust output of an afterburning turbojet engine at the higher 
Mach numbers encountered with supersonic aircraft. For example, the after- 
burning net thrust obtained with a simple convergent nozzle at Mach 1.7 can 
be increased by approximately 10 per cent when a properly designed con- 
vergent-divergent jet nozzle is used. This increase in thrust over that obtained 
by the use of a simple convergent nozzle can be acquired with no increase in 
turbine temperature or engine throttle setting and with little increase in engine 
weight and nozzle complexity. Clearly, the convergent-divergent type of nozzle 
is both a desirable and practical component for supersonic aircraft. 

Thrust Reversers 

The difficult problem of stop- 
ping an aircraft after landing 
increases manyfold with higher 
airspeeds and greater gross 
weights which result in higher 
Figure 2—30. Thrust Reverser in Operation wing loadings and increased 

landing speeds. In many in- 
stances, wheel brakes can no longer be entirely relied upon to slow the 
aircraft within a reasonable distance, immediately after touchdown. The re- 
versible pitch propeller has solved the problem for reciprocating engine air- 
craft and turboprop-powered airplanes. Turbojet aircraft, however, must rely 





PWA OPER. | NSTR. 200 

upon either some device such as a parabrake or runway arrester gear or some 
means of reversing the thrust produced by their engines. 

Although in general use on military aircraft, the parabrake or drag parachute 
has distinct disadvantages and probably will never prove practical for com- 
mercial aircraft. The parabrake is always subject to either a premature open- 
ing or failure to open at all. The parabrake must be recovered and repacked 
after each use or, if damaged or lost, must be repaired or replaced. Once the 
parabrake has opened, the pilot has no control over the amount of drag on 
the aircraft except to release the parachute completely. Arrester gears are 
primarily for aircraft-carrier deck operation and may occasionally be used as 
land runway overshoot barriers. They would hardly be suitable for normal 
airport runway operation. 

An engine thust reverser (Figure 2-30), on the other hand, not only provides 
a ground-speed braking force but, if suitable, is desirable for use prior to 
landing. The requirement for operating a turbojet engine at high rpm during 
final approach, in order to minimize the length of time needed to accelerate 
the engine in the event of a “go-around,” can be met without unduly increas- 
ing the airspeed of the aircraft, when a thrust reverser is employed. The use of 
a thrust reverser with partial, as well as full, reverse thrust capabilities will 
enable the approach to be made not only at full rpm but also with the effective 
forward thrust reduced to any desired amount. Rapid descent from altitude 
can be accomplished more effectively and without compromising the perform- 
ance of airbleed-driven accessories or anti-icing equipment. Chances of a 
flameout are minimized as well. 

At present, turbojet engine thrust reversers are still in the experimental and 
testing stage. Quite a number of devices for reversing thrust have been pro- 
posed, many of which have been tested with a considerable degree of success. 
Thrust reversers will become standard equipment in the very near future on 
commercial airline and military transport-type aircraft equipped with nonafter- 
burning engines. It is also entirely possible that military fighter-type aircraft 
of the more distant future will employ a reverse thrust device to replace today’s 
aircraft dive or speed brake. 

The most successful thrust reversers can be divided into two categories, the 
mechanical blockage type and the aerodynamic blockage type. Mechanical 
blockage is accomplished by placing a removable obstruction in the exhaust 
gas stream, usually somewhat to the rear of the nozzle. The engine exhaust 
gases are mechanically blocked and diverted at a suitable angle in the reverse 
direction by means of an inverted cone, half-sphere or other means of obstruc- 
tion which is placed in position by the pilot to reverse the flow of exhaust 
gases. In the aerodynamic blockage type of thrust reverser, thin airfoils or 
obstructions are placed in the gas stream, either along the length of the exhaust 
duct or immediately aft of the exhaust nozzle. In one adaptation of the aero- 
dynamic reverser, vanes inside the duct create swirling of the gases in a man- 
ner which centrifuges them into a cascade of turning vanes. In another form, 
a pneumatic or mechanically operated obstacle is introduced into the gas 
stream to cause lateral deviation of the gas flow into a series of turning vanes. 





An acceptable thrust reverser must not affect engine operation either when 
the reverser is operating or when it is not. It must be able to withstand high 
temperatures, and must be mechanically strong, relatively light in weight, 
reliable and “fail-safe.” When not in use, it should not add appreciably to 
the engine frontal area and must be streamlined into the configuration of the 
engine nacelle. In order to satisfy the minimum braking requirements after 
landing, a thrust reverser should be able to produce in reverse at least 50 
per cent of the full forward thrust of which the engine is capable. 

The clamshell type of mechanical blockage reverser appears to satisfy all of 
these requirements best and, in one form or another, will undoubtedly be the 

one adopted for general nonafterburner engine: 

use. The W-clamshell shown in Figure 2-31 
has operated very satisfactorily. When open, 
the reverser is approximately one nozzle diam- 
, eter to the rear of the engine exhaust nozzle. 

Open Retracted When folded and not in use, it nests neatly 

around the engine exhaust duct, forming the 

Figure 2—31. W-Clamshell Tear section of the engine nacelle cowling. The 

Mechanical Blockage Thrust W-clamshell reverser has the added advantage 

Reverser of splitting the exhaust gas discharge, which 

tends to prevent flow instability and severe 

buffeting. The clamshell is opened and closed by pneumatic pistons operated 

by high-pressure bleed air from the engine compressor. The speed at which 
the reverser operates is sufficient to meet most emergencies. 

Exhaust Silencers 

es ul nun _| 

Commercial turbojet aircraft operating in and out of airports located in or 
near thickly populated areas will unquestionably require that some sort 
of silencing device or noise suppressor be provided for the engine, particularly 
during the take-off and initial climb phase of flight. It is desirable to moderate 
the noise produced by a turbojet engine to a level which will be no more 
objectionable than that of a reciprocating engine and propeller, operating 
under similar conditions. The manner in which the noise of a turbojet aircraft 
can be reduced to a level as acceptable as that of a reciprocating engine air- 
craft is not simple to determine. The propeller, which is a major source of 
noise in reciprocating engine aircraft, has a noise pattern which rises sharply 
to a maximum level as the plane of the propeller passes an individual on the 
ground and then drops off almost as sharply after the propeller has gone by. 
The turbojet aircraft produces a sharp rise in noise which reaches a peak after 
the aircraft has passed an individual on the ground and is at an angle of 
approximately 45 degrees to him. The noise then persists at a high level for a 


r PWA OPER. [NSTR. 200 

considerable period of time as compared with that of a reciprocating engine 
fa with a propeller (Figure 2-32). 




-60 0° CO a 60 100 140 

Figure 2—32. Noise Level, with Average Background 



There are three sources of noise involved in the operation of a gas turbine 
engine. The engine air-intake and vibration from the engine housing are 
sources of some noise but the noises thus generated do not compare in magni- 
tude with that produced by the engine exhaust. The noise produced by the en- 
gine exhaust is caused by the high degree of turbulence of a high-velocity jet 
stream moving through a relatively quiescent atmosphere (Figure 2-33). Fora 

| aOR D Soe eae aaa eget 
po ae 

i ao oa) 

Most of the nose radiates 
from this low frequency 
turbulence region 
D=Nozzle Diameter 

Figure 2—33. Turbojet Exhaust Noise Pattern 

distance of a few nozzle diameters downstream behind the engine, the velocity 
of the jet stream is high and there is little mixing of the atmosphere with the 






jet stream. In this region, the turbulence within the high-speed jet stream is 
very fine-grain turbulence and produces relatively high-frequency noise. 

Farther downstream, as the velocity of the jet stream slows down, the jet 
stream mixes with the atmosphere and turbulence of a coarser type begins. 
Noise from this portion of the jet stream has a much lower frequency. As the 
energy of the jet stream finally is dissipated in large turbulent swirls, a greater 
portion of the energy is converted into noise. The noise generated as the 
exhaust gases dissipate is at a frequency near the low end of the audible range. 
The lower the frequency of the noise, the greater the distance that it will 
travel. This means that the low-frequency noises will reach an individual on 
the ground in greater volume than the high-frequency noises and hence will be 
more objectionable. High-frequency noise is weakened more rapidly than low- 
frequency noise, both by distance and the interference of buildings, terrain 
and atmospheric disturbances. A deep-voiced, low-frequency foghorn, for ex- 
ample, may be heard much farther than a shrill high-frequency whistle, even 
though both may have the same over-all volume (decibels) at their source. 
Noise levels vary with engine thrust 
(Figure 2-34) and are proportional 
to the amount of work done by the 
engine on the air which passes 
through it. An engine having rela- 
tively low airflow but high thrust 
due to high turbine discharge (ex- 
haust gas) temperature, pressure 
and/or afterburning will produce 
a gas stream of high velocity and 
therefore high noise levels. A larger 
engine, handling more air, will be 



quieter at the same thrust. It fol- Figure 2—34. Noise Level vs Engine Thrust 

lows, then, that the noise level can 

be considerably reduced by operating the engine at lower throttle settings 
and that large engines operating at partial thrust will be less noisy than 
smaller engines operating at full thrust. 

Because of the characteristic of low-frequency noise to linger at a relatively 
high volume, effective noise reduction for a turbojet aircraft must be achieved 
either by revising the noise pattern in such a manner that the high noise level 
comes and goes quickly as the aircraft passes by, as in the case of a propeller- 
driven reciprocating engine aircraft, or by reducing the peak noise level to a 
volume appreciably lower than that produced by propeller-driven aircraft if 
it is to be equally acceptable to an individual on the ground. The ultimate 
answer is probably a combination of both methods. Deep grooves or a multi- 
tude of small nozzles around the engine nozzle, itself, tend to reduce the level 
of noise in the lower frequencies by introducing turbulence in the jet stream 
near the nozzle. This has the effect of altering the noise pattern in a manner 
which reduces the time that it takes the peak noise to pass a given point. 

Since silencers of sufficient value for general use are still in the development 
Stage, it is impossible to describe them or to explain the method by which they 


r PWA OPER. [NSTR. 200 
r i. will operate. Regardless of the ultimate form of silencers, the basic problem is 

and will continue to be one of balancing results with the penalties involved. 
r It is a question of how much additional weight and loss of thrust can be 

tolerated in order to achieve a specific amount of reduction in the noise pro- 

duced by a turbojet engine. That some penalty will be necessary is axiomatic. 
P It remains for the development engineer to determine how much noise reduc- 
| tion can be achieved and for the aircraft operator to compromise between the 

added cost involved and the demands of the general public living in the 
f immediate vicinity of operating airports and air bases. 

P Afterburners 

Afterburning is a method by which the maximum thrust capability of a basic 
engine may be augmented by an additional 50 per cent or more. In slang 
P phraseology, the use of an afterburner is said to be “hot” operation. When the 
afterburner is off, the term “cold” operation is used. Afterburning is known 
as “reheating” in Great Britain. 

oo | There are occasions when more than the 
: full thrust of an engine is required for 
| short periods of time, such as to reduce 
the length of the take-off run, to increase 
Fr the rate of climb or to provide an extra 
burst of speed during a combat intercept 
mission. It would not be economical to in- 
P stall a larger engine and thus penalize the 
aircraft with added engine frontal area, 
weight and specific fuel consumption just 
f to satisfy a periodic need for more thrust. 
The solution is an afterburner. It has been pointed out (Section I) that an 
- afterburner takes advantage of the fact that approximately 25 per cent of the 

air passing through a turbojet engine is consumed by combustion, the remain- 
ing 75 per cent being capable of supporting additional combustion if more 
P fuel is added. 

Fundamentally, an afterburner is a ramjet engine attached to the exhaust of 
a turbojet engine. The high air velocity initially necessary to enable the ramjet 
to operate is supplied by attaching the afterburner to the turbine discharge 
section of a conventional turbojet engine. The gases leaving the basic engine 
F | have sufficient velocity at the higher thrust settings to satisfy ramjet require- 
| ments amply, regardless of whether the aircraft is in a steep dive or standing 

at rest at the end of a runway, ready for take-off. The remarkable thing about 
a on an afterburner is its simplicity. It consists of only four fundamental parts: the 

afterburner duct, the fuel nozzles or spraybars, the flame holders, and a two- 


PWA OPER. | NSTR. 200 


position or variable-area exhaust nozzle. When the afterburner is not operat- 
ing, the afterburner duct serves as the basic engine tailpipe. 



The afterburner duct is the main element of an afterburner. It must be shaped 
in such a manner that it does not upset the operating pressures of the basic 
engine. With a properly designed afterburner, the engine, itself, is completely 
unaware of the afterburner’s existence. The engine operates essentially as 
usual, whether the afterburner is operating or not. The afterburner duct must 
be of such proportions that stable combustion can be maintained during after- 
burning operation. This requires a burning section of sufficient cross-sectional 
area to ensure that the gas velocity through the afterburner does not exceed 
the rate of flame propagation. Otherwise, the flame would not be able to 
establish a firm foothold because the onrushing turbine exhaust gases would 
push the burning mixture right out of the exhaust nozzle. Fuel is introduced 
through a series of simple perforated spraybars located inside the forward 
section of the afterburner duct. Not far aft of these, flame holders are pro- 
vided to help create local turbulence and to reduce the gas velocity in the 
vicinity of the flame. The flame holders may take the form of concentric 
rings of an angular “V” cross section or they may be of a cross-sectional 
“C” or “U” shape. | 

When the afterburner is in operation, fuel is introduced through the fuel 
nozzles and is burned within the duct, causing considerable additional expan- 
sion of the gases after they leave the basic engine. This increases their momen- 
tum and provides more thrust. When an afterburner is used, the increase in 
total fuel consumption is approximately threefold. However, this sizable 
increase in specific fuel consumption is more than offset by the results obtained 
in added aircraft performance for special-purpose use, such as a faster climb. 
Afterburners are normally found only on specific types of military aircraft, 
as the added weight would not be justified for only periodic use during take-off 
or initial climb on transport or commercial aircraft where economy and noise 
are more important than occasional high performance. 

The afterburner is ignited by introducing a flash of flame through the turbines. 
The flash of flame is created by Squirting a stream of raw fuel into the burner 
section of the basic engine. This is normally accomplished automatically when 
the afterburner is turned on. The afterburner igniter and the afterburner fuel 
control are considered parts of the engine fuel system and are discussed later. 
As more fuel is metered by the fuel control to the afterburner, combustion 
temperature is increased, the reheated engine exhaust gases expand more and 
more, additional momentum is imparted to the exhaust stream as it accelerates 
rearward, and increasing amounts of thrust augmentation are produced. The 
degrce of augmentation is regulated by varying the amount of fuel to be mixed 
with the unburned air from the basic engine exhaust. 




Afterburner Exhaust Nozzle 

Because of the increase in expansion ratio of the gases within the afterburner 

when it is in operation, the afterburner exhaust nozzle must be capable of 
being increased in area during afterburning in order to assure that basic engine 
operation will continue to be normal. The exhaust nozzle also must be capable 
of being closed to its original small area during nonafterburning (cold) opera- 
tion. An adjustable nozzle is provided which may be of the clamshell type 
(Figure 2-35). Sometimes an iris arrangement is used. A series of flaps around 

Figure 2—35. Two-position Adjustable Afterburner Nozzle 

the perimeter of the nozzle, somewhat similar in design to reciprocating engine 
cowl flaps (only, of course, much heavier and stronger), is yet another type 
of adjustable nozzle design. Opening or closing the variable-area exhaust 
nozzle controls the expansion ratio of the gases through the afterburner. 
During cold (nonafterburning) operation, the exhaust nozzle offers the small- 
est amount of area in order to maintain the normal expansion or pressure 
ratio of the gases within the basic engine. As the afterburner commences to 
function (hot operation), the nozzle area is increased so that the proper ratio 
‘s maintained. Nozzles having areas which may be varied throughout the 
entire range of operation are in the offing for the future. They will be adjusted, 
probably automatically, to coincide with the amount of thrust called for by 
the cockpit throttle position. Although an afterburner usually may be op- 
erated at either full or partial afterburning thrust, present-day nozzles are 
fully open whenever the afterburner is on. 

Afterburner Exhaust Nozzle Control 

At present, the most generally used type of afterburner exhaust nozzle control, 
or “ENC” as it is frequently called, operates by sensing fuel pressure in the 
afterburner fuel system. Engine internal pressure, tapped at the diffuser case, 
is used to move a piston in the afterburner nozzle actuator to open and close 
the afterburner exhaust nozzle automatically. When the afterburner is turned 




on, the pressure of metered fuel in the afterburner fuel system, working in 
conjunction with a spring, overcomes the pressure of unmetered fuel on the 
opposite side of a transfer valve. This moves the transfer valve to a position 
which shunts the engine internal pressure to the side of the nozzle actuator 
piston which will open the afterburner exhaust nozzle. When the afterburner 
is turned off, the pressure of unmetered fuel, alone, working against the spring, 
is sufficient to overcome the spring tension and cause the transfer valve to 
shunt engine internal pressure to the side of the nozzle actuator piston which 
will close the afterburner exhaust nozzle. 

A more advanced type of afterburner exhaust nozzle control (ENC) operates 
entirely by engine internal pressure, thus eliminating any possible fire hazard 
which might develop should a fuel leak occur in the ENC or in the fuel line 
to it. In this type of nozzle control, turbine discharge pressure is piped to the 
ENC to take the place of the unmetered fuel pressure used in the fuel pressure- 
operated type of control. As will be shown later, the afterburner igniter 
operates by means of engine internal pressure, tapped from the afterburner 
nozzle actuator open-position pressure line. When the afterburner igniter is 
actuated, engine internal pressure is shunted to the ENC to oppose and over- 
come turbine discharge pressure, just as metered fuel pressure Opposes and 
Overcomies unmetered fuel pressure in the older type of ENC when the after- 
burner is turned on. When the afterburner is turned off, turbine discharge 
pressure, alone, is sufficient to overcome the tension of the spring in the ENC 
and cause the nozzle to close. 

Another innovation in afterburner exhaust nozzle control, which may oc- 
casionally be a feature of some engines, is known as the “pop-open” nozzle. 
On engines equipped with a nozzle of this type, an exhaust nozzle control 
override, or some similar device, causes the afterburner exhaust nozzle to 
open when the throttle is retarded far enough to encounter the “Idle flat” 
on the fuel control metering cam during engine deceleration. During accelera- 
tion, the nozzle closes again as soon as the throttle is advanced above the 
Idle flat. Thus, during nonafterburning engine operation, the afterburner 
nozzle is in the open position only when the throttle is fully retarded. The 
open nozzle serves to “dump” engine thrust at Idle, permitting the Idle rpm 
to be increased without exceeding the thrust at Idle for which the engine 
was designed. The pop-open nozzle serves to provide more bleed air for the 
operation of accessories when the engine is idling and permits more rapid 
engine response when the throttle is advanced. Better response of the engine 
during acceleration from Idle makes the engine more flexible for operation 
from an aircraft carrier. The obvious disadvantages of a nozzle of this type 
are its complexity, with the accompanying increased Opportunity for mal- 
function, and the added weight involved. 

Afterburner Screech Liners 

Afterburners are occasionally subject to a type of combustion instability 
known as “screech.” Screech is a condition of periodic violent pressure fluctua- 
tions in the afterburner duct resulting from cyclic vibration due to unsteady 
release of combustion energy. Screech is characterized by intense noise and 
frequently results in structural failure of the afterburner duct. When screech 






occurs, heat transfer rates and temperatures of the afterburner parts increase 
greatly. Moderate to severe screech can cause rapid deterioration or failure of 
the flame holders or the afterburner duct. Screech is controlled by placing 
so-called “screech liners” in the duct. These are inner steel sleeves which line 
the inside of the duct. The special design of the sleeves tends to absorb the 
periodic combustion energy fluctuations and to prevent random pressure 
fluctuations from developing into cyclic vibrations of large amplitude. 


Accessories for gas turbine engines can be divided into two categories: those 
driven by high-pressure bleed air taken from the rear of the engine compres- 
sor, and those driven mechanically by an accessory drive shaft and gear box 
connected directly to the compressor shaft. The compressor interstage bleed 
air which is ducted overboard to obviate compressor stall during certain 
periods of low-thrust operation is not used for this purpose since it is not avail- 
able at high thrust settings nor is it steady in volume or pressure. 

Airbleed-driven Accessories 

Gas turbines are unique among 
engines in that high-pressure air is 
readily available for driving air- 
craft accessories by means of air 
turbine motors which operate in- 
Figure 2—36. Axial Compressor dependently of the engine, itself. 
Accessory Drive Airbleed Ports Compressor discharge air at high 
pressure and temperature is bled 
from the engine through ports which are usually located in the diffuser case 
(Figure 2-36). The pressure, of course, is the result of the compression 
achieved by the compressor. The temperature is the result of the work done 
on the air during the compression process. The air is ducted as a source of 
power for operating such items as air-conditioning units, hydraulic pumps, 
generators, and the like. Cockpit or cabin pressurizing and heating units 
sometimes are operated in this manner by means of a separate airbleed-driven 
compressor, although, in some installations, compressor bleed air is piped 
directly to the cockpit for the purpose. The first method guarantees pure un- 
contaminated air for human consumption, but is somewhat expensive in terms 
of weight and power required. Although more economical, cockpit or cabin 
pressurization by use of direct compressor airbleed is usually restricted to 
military aircraft in which oxygen masks are available to all crew members 
should something go amiss and the air become contaminated with minute oil 
particles, smoke or some other substance coming from the interior of the en- 
gine. On multiengine aircraft equipped with pneumatic engine starters, one 
engine is usually started from a ground air source, then air from this oper- 
ating engine is bled through a system of ducts in the aircraft to be used to 
turn the starters of the other engines. Because of its high temperature, bleed 
air is also used for windshield and wing deicing and for compressor air- 
inlet anti-icing. 




The amount of air available for driving accessories and for other purposes in 
the aircraft is usually from 1 to 3 or 4 per cent of the total airflow through 
the engine. It must be borne in mind that the air under pressure which is 
extracted from the engine is not a bonus. Power extraction is only obtained 
at a sacrifice either in engine output or fuel consumption. 

Mechanically Driven Accessories 

The other method of driving acces- 
sories is by means of a mechanical, 
geared drive taken directly from 
the main shaft connecting the tur- 
bine to the compressor (Figure 
2-37). Accessory drives and acces- 
Figure 2—37. Axial Compressor sory mounting pads are provided 
Mechanical Accessory Drive by the engine manufacturer. These 
are located in the nose section of 
the engine or, sometimes, on the side of the engine in the vicinity of the rear 
of the compressor section. Dual axial compressor engines usually have two 
accessory drive gear boxes, one being connected to the low pressure compres- 
sor and the other to the high pressure compressor. This method of driving 
accessories is used for tachometers, hydraulic pumps, generators, alternators, 
and other units which can or must be mounted close to the engine or which 
must be connected directly to the engine, such as a tachometer. 

= Se is 


Gas turbine engines are started by rotating the compressor. In the case of dual 
axial compressor engines, the high compressor is usually the only one rotated. 
First, it is necessary to accelerate the compressor to provide sufficient air, 
under pressure, to support combustion in the burners. Secondly, once fuel has 
been introduced and the engine has fired, the starter must continue to assist 
the engine above the self-sustaining speed of the engine. The torque supplied 
by the starter must be in excess of the torque required to overcome compressor 
inertia and the friction loads of the engine. 

The basic types of starters which have been developed for gas turbine engines 
are d-c electric motor, air turbine and combustion. An impingement starting 
system is sometimes used on small engines. An impingement starter consists 
of simple jets of compressed air piped to the inside of the compressor or tur- 
bine case in such a manner that the jet air blast is directed onto the compressor 
or turbine rotor blades, causing them to rotate. Electric motors are chiefly 
used in engine ground test stands where bulk and weight are not important 
and where ample electric power is readily available. 

Figure 2-38 graphically illustrates a typical starting sequence for a gas turbine 
engine. As soon as the starter has accelerated the compressor sufficiently to 




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establish airflow through the en- 
gine, the ignition is turned on, and 
then the fuel. The exact sequence 
of the starting procedure is im- 
portant because there must be suf- 
ficient airflow through the engine 
to preclude the danger of an ex- 
plosion before the fuel/air mix- 
ture is ignited. The fuel flow rate 
will not be sufficient to enable the 
engine to accelerate until after self- 
sustaining speed has been attained. 
If assistance from the starter were 
cut off below the self-sustaining 
speed, the engine would decelerate 
because it cannot produce enough 
energy to sustain rotation during 
the initial phase of the starting 
cycle. The starter must therefore 
continue to assist the engine CON- C€tcrter'on” ‘Time — Seconds 

siderably above the self-sustaining 

speed to avoid a delay in the start- — Figure 2—38. Typical Starting Sequence 
ing cycle which would result in a for a Gas Turbine Engine 

hot or hung (false) start. At the 

proper points in the sequence, the starter and, usually, the ignition will be 
automatically cut off. The higher the rpm before the starter cuts out, the 
shorter will be the total time required for the engine to attain Idle rpm 
because the engine and the starter are working together to furnish torque 
above the self-sustaining speed. 


Compressor- RPM ——_> 

Exhaust Gas Temperature — °C 


Air Turbine Starters 

Air turbine or pneumatic starters are probably the most common type em- 
ployed on gas turbine engines. A small, geared air turbine is attached to the 
engine starter pad located at the accessory drive section. The air turbine serves 
to accelerate the compressor. Air is supplied from a ground cart which may 
contain a small, shaft turbine engine. This engine can supply air from its own 
compressor or it may drive a separate compressor. On multiengine aircraft, 
air is often bled from the first engine started and is used to operate the starters 
of the remaining engines. With an air turbine (pneumatic) starter, it is im- 
portant that the air supply be of sufficient volume and pressure to meet starter 
specifications. Otherwise, the starter torque may not be adequate to produce 
consistently successful starts within an acceptable limit of time. 

aa ee a 

Combustion Starters 

A combustion starter is essentially a small turboshaft gas turbine engine. 
Its chief advantages are quick starts, because of the high torque produced, 


So = sii ac a ts 




————— oe 



and the fact that the complete starting system may be carried aboard the air- 
craft. Should a landing be made where a suitable outside high-pressure air 
supply for starting is not available, the engines may be started with the air- 
borne equipment. This type of starter obtains its power from hot expanding 
gases generated within the combustion chamber of the starter by burning a 
combustible mixture which may be either fuel and air or a specially com- 
pounded solid or liquid slow-burning substance called a monopropellant. 
The fuel/air type of starter requires a source of fuel and high-pressure air, 
the air usually being furnished from a compressed air source on the ground 
or an air bottle carried aboard the aircraft. The fuel may be either from the 
aircraft fuel tanks or from a special tank carried specifically for the purpose. 
The monopropellant type of starter does not need air to Support combustion. 
Therefore, no external source of fuel and air is necessary. In some starters, the 
monopropellant may be inserted in cartridge form, as needed. Both types of 
combustion starter are electrically ignited and begin to operate automatically 
as soon as the starter switch is closed. 

The quantity of fuel and air or monopropellant required by the starter is 
directly proportional to the length of time that the starter is operating. Weight 
and size limitations of aircraft necessitate that the quantity of fuel and air or 
monopropellant be held to a minimum in an airborne combustion starter 
system. It is therefore essential that the starter burning time be held to a 
minimum. The minimum starter burning time is determined by the starter 
torque and the starter cutout speed. The starter cutout speed depends upon 
the self-sustaining speed of the engine. The self-sustaining speed, then, deter- 
mines, to a large extent, the amount of fuel and air or monopropellant re- 
quired. Ordinarily, sufficient fuel and air or monopropellant will be carried 
aboard the aircraft to provide a minimum of two starts. Most combustion 
starters can also be made to operate as simple air turbines. When an outside 
air source, either from a ground cart or from another engine which is already 
started, is connected to the starter, the combustion starter will function as a 
pneumatic starter. 

Fuel Systems and Fuel Controls 

The primary function of an aircraft fuel system is to supply clean fuel, free 
from vapor, at the proper pressures and flow rates to the engine under all 
operating conditions. In general, the system must be designed, and its com- 
ponent parts selected, to satisfy the particular requirements of the aircraft. 
Turbine-powered, rapid-climbing, high-altitude aircraft present difficult prob- 
lems to the designer of a satisfactory fuel system. High rates of fuel flow, 
low atmospheric pressure in the fuel tanks, complexity of piping systems, 
high weight and large size of the engine installation, vapor loss with con- 
sequent reductions in range, and cold-weather starting are all factors which 
must be considered. 

Figure 2-39 illustrates a typical fuel system for a nonafterburning engine 
equipped for water injection. Figure 2-40 is a typical fuel system for an after- 
burning engine. A study of these diagrams will show that the fuel system con- 



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Figure 2—39. Fuel System, Nonafterburning Engine Equipped For Water Injection 

sists of basic essential elements to which a water injection system is added in 
one case, and an afterburner system in the other. For some engines, the basic 
system may be used alone. 

The Basic Fuel System 

The system described here is more or less typical for a dual axial compressor 
turbojet engine. Although other fuel systems may differ somewhat, the same 




elements, performing essentially the same functions, usually will be present in 
systems for other types of gas turbine engines. 

Fuel Pump 

The fuel pump is a separate unit placed just ahead of the main fuel control. 
Fuel is supplied from the aircraft fuel tanks through the necessary filters, boost 
pump and valves to the engine-driven fuel pump. This is usually a multistage, 
spur-gear-type, positive-displacement, high-pressure pump. Fuel is delivered 
from the pump to the fuel control where it is metered to the engine in proper 
quantities. Excess fuel is bypassed back to the pump. 

Emergency Fuel System 

Because of possibly serious consequences of fuel flow interruption in single- 
engine airplanes, fighter-type aircraft engines usually are equipped with fuel 
controls in which an emergency fuel system is incorporated. The emergency 
fuel system is sometimes called the manual system. The emergency section 
of the fuel control parallels the primary or normal fuel system. An electrical 
switch in the cockpit enables the pilot to select manually either the normal or 
emergency system of the fuel control. The emergency system is simpler than 
the normal system and frequently lacks the acceleration-limiting and rpm- 
governing capabilities of the normal system. Fuel flow regulation must be 
accomplished by manual throttling. Turbine inlet temperature in the engine 
is controlled solely by throttle position and by monitoring the exhaust gas 
temperature gage visually. The emergency fuel system permits sustained opera- 
tion and may be used for engine starting and afterburning, but requires close 
cockpit attention to assure that critical limits are not exceeded. Desirable 
though it is, an emergency or manual fuel system represents added weight and 
complication, which usually preclude its use on multiengine aircraft in which 
an engine failure will not be particularly hazardous. 

Pressurizing and Dump Valve 

From the fuel control, the fuel flows through a flowmeter to a pressurizing 
and dump valve, passing on to the engine through either the primary fuel 
manifold, alone, or, at the higher fuel flows, through both the primary and 
secondary manifold. These sometimes are called the pilot and main manifold, 
respectively. Of the two, the primary (pilot) manifold is the smaller in size. 
Both manifolds spray fuel into the engine burners through a system of nozzles. 

At the beginning of an engine start, the fuel control supplies a pressure signal 
to the pressurizing and dump valve, causing the valve to close the manifold 
drain and to open a passage for fuel flow to the engine. This signal keeps the 
valve open during all engine operation. On engine shutdown, fuel flow is cut 
off immediately by the throttle in the cockpit. The strength of the pressure 
signal to the pressurizing and dump valve drops rapidly as the metered fuel is 
directed back to the fuel pump inlet instead of to the engine. As soon as the 
pressure signal to the valve becomes inadequate to hold the valve open, the 
passage for fuel flow to the engine will close, permitting the fuel manifolds to 
drain overboard. 



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Figure 2—40. Fuel System, Afterburning Engine 

Afterburner Fuel System and Fuel Control 

In afterburning engines, a stage of the engine-driven fuel pump supplies fuel 
to another fuel system in addition to that of the main fuel control. A fuel 
transfer valve, mounted on the body of the fuel pump and operated by a 
motor, directs fuel either back to the afterburner pump stage inlet or to an 
afterburner fuel control. The afterburner fuel control senses the pressure at 
the diffuser case and meters fuel to the afterburner fuel nozzles or spraybars 




accordingly. Excess fuel is returned to the pump. A drain valve which is 
usually provided in the afterburner manifold will be closed while fuel is flow- 
ing in the afterburner manifold but will open to drain the fuel in the manifold 
overboard on engine shutdown. 

Afterburner Igniter 

For purposes of starting the afterburner, an igniter receives metered fuel from 
the afterburner manifold. Pressure, tapped from the afterburner nozzle actu- 
ator open-position line, is used to squirt a stream of fuel into one of the 
engine burner cans. This causes a flash of flame to pass through the turbines 
to the point where the afterburner nozzles are delivering fuel. If the after- 
burner fails to light, the igniter must be relieved of fuel pressure by stopping 
the afterburner fuel flow (by retarding the throttle), following which another 
light may be attempted. 

A “circulating” afterburner igniter is essentially the same as a conventional 
igniter except that, rather than filling only the fuel charge chamber of the 
igniter when an afterburner light is called for, fuel circulates continually 
through the chamber keeping it full, ready for a light at all times. This makes 
for quicker operation of the afterburner igniter and usually more consistently 
successful afterburner lights at high altitude. 

Afterburner igniter systems may be either of the open or closed type. In the 
more conventional open system, the igniter operates after the afterburner 
exhaust nozzle has opened. In the so-called “nozzle closed light” system, the 
afterburner igniter operates just prior to the opening of the nozzle. Although 
the open system is probably the more foolproof of the two because it offers 
less chance of afterburner operation with the nozzle in closed position, the 
closed light system is intended to provide better afterburner lighting charac- 
teristics at high altitude. 

Water Injection and the Fuel System 

The fuel control plays a major part in the water injection process on non- 
afterburner engines. A switch, incorporated in the fuel control and operated 
by the throttle, provides electrical power for a motor-actuated, water-shutoff 
valve. When water is flowing, a reset signal is fed back to the fuel control 
linkage from the water regulator, which adjusts the metering of fuel to the 
engine in a manner suitable for water injection engine operation. 

Water, supplied from the aircraft water-storage tank by a pump, passes 
through strainers and valves to the water-shutoff valve. When the throttle is 
placed in a predetermined position and the injection system is armed by means 
of an arming switch, the motor-actuator opens the shutoff valve, permitting 
water to flow to the water regulator. The regulator provides a constant pres- 
sure head to an orifice which, in turn, provides a constant water flow to the 
engine. The water then passes on to the engine water inlets where it is injected 
through a spray ring. 



— ca 

P -~ Fuel Controls 

Strictly speaking, the pilot of a turbojet or turboprop aircraft does not control 
his engine. He acts through an intermediary, the fuel control. His relation to 
his powerplant corresponds to that of the bridge officer on a ship who obtains 
engine response by conveying his orders to an intermediary, the engineering 
officer, who, in turn, translates the directions from the bridge into appropriate 
action. In both cases, the intermediary monitors certain factors which are not 
discernible to the top command, and, when necessary, may take or withhold 
action without waiting for topside instructions. 

The gas turbine counterpart of the ship’s engine-order telegraph is the pilot’s 
throttle. This is*not a throttle in the reciprocating-engine sense of the word 
because the turbojet throttle in the cockpit does not give the pilot direct con- 
trol over the fuel-throttle valve in the fuel control. By placing the turbojet 
cockpit throttle in a given position, the pilot tells the fuel control how much 
thrust he desires in approximate percentages of full thrust. The control mon- 
itors certain variables and provides sufficient fuel flow to the engine to pro- 
duce the desired thrust (or power, if the engine is a turboprop), but at a 
P flow rate which will not allow the engine operating limits to be exceeded. 

=> Ge a 

= we 

Turbojet and turboprop engines are equipped with a control arm on the fuel 
control to which the throttle lever in the aircraft is connected through a sys- 
tem of pulleys and cables, or other appropriate linkage. On some engines, 
eo this may be an engine cross shaft, the throttle linkage being connected to an 
arm on one end of the shaft. The other end of the shaft is connected to the 
fuel control. Whether it be a control arm on the fuel control or an arm on the 
engine cross shaft, it is this control arm to which reference is made whenever 
the term, “power lever,” appears in various Pratt & Whitney Aircraft publica- 
tions. The reference to “throttle” specifically means the “go handle” in the 
- cockpit of single-engine aircraft or on the control pedestal and the engineer’s 
panel of multiengine aircraft. 

a “= 

Fuel controls for many centrifugal and single axial compressor engines vary 
the thrust by controlling rpm. In engines of these types, rpm and thrust are 
proportional when turbine inlet temperature is allowed to vary inversely with 
compressor inlet temperature. It’s a different story in the case of a dual axial 
compressor engine. Dual axial compressor engines must be protected when 
engine thrust increases on cold days. This is done at some expense of thrust 



F on hot days when engine thrust is reduced. However, to get all of the thrust 
possible on hot days and still protect the engine on cold days, it is neces- 
- sary to control turbine inlet temperature to constant values and allow 

rpm to vary. This is the task of the fuel control. Although all fuel controls 
include many of the same basic features, the discussions which follow pertain 
r specifically to a type of control which might be found on a dual axial com- 
pressor engine, such as a Pratt & Whitney Aircraft J57 for military use. Fuel 
controls for Pratt & Whitney Aircraft commercial engines of this type may 
rm be somewhat simplified (depending upon the desires of the airline operator) 
by elimination of some of the automatic sensing devices. 




The fuel control senses compressor inlet temperature as a function of the 
density of the air entering the engine. It also senses high-pressure compressor 
rpm and engine burner pressure. These three variables affect the amount of 
thrust an engine will produce for a given fuel flow. Because compressor inlet 
pressure is also a function of atmospheric density, inlet pressure is sensed by 
some controls instead of inlet temperature. However, most inlet-pressure 
sensing controls adjust the acceleration schedule by additionally sensing inlet 
temperature. By taking cognizance of burner pressure, the control can also 
limit fuel flow so that allowable engine internal pressures will not be exceeded, 
particularly on a cold day at low altitude, such as during take-off. Upon receipt 
of a signal from the pilot for a given level of thrust, the control takes all of the 
monitored variables, including rpm, into consideration. Then, as may be 
necessary, the control increases or decreases fuel flow to provide the required 
thrust. As long as the throttle in the cockpit remains in a given position, the 
control will vary fuel flow with changing compressor inlet conditions to main- 

tain the approximate percentage of full engine thrust called for by the throttle 

setting. Turbine inlet temperature for any given increment of thrust will 
remain approximately constant. 

Because turbine inlet temperature is the ultimate limiting factor of gas turbine 
engine operation, the thrust ratings of dual axial compressor engines are based 
upon the allowable turbine inlet temperatures, either for a time-limited period 
or for continuous operation. The Military thrust rating, or Take-off (dry) 
rating for commercial engines, is the maximum thrust which the engine will 
produce without exceeding the maximum allowable turbine inlet temperature 
which may be used for a specified limited period of time. Normal Rated (mili- 
tary engines) or Maximum Continuous (commercial engines) is the thrust 
produced at the maximum turbine inlet temperature which may be used 
continuously. Centrifugal and single axial compressor engines often are rated 
according to the allowable compressor rpm. 

In the interest of obtaining thermal efficiency and maximum thrust, the tur- 
bine inlet temperature is maintained as close to the allowable limits as possi- 
ble. However, this temperature is not measured by most fuel controls. 
Measuring turbine inlet temperature is a liability for the turbines. A satis- 
factorily sensitive measuring probe might break and pass through the turbine 
wheels. One which was sufficiently strong would be too sluggish. The problem 
is resolved by calculating the turbine inlet temperatures which will be pro- 
duced under various conditions. The predetermined schedule is calculated to 
ensure turbine operation within the desired or safe temperature limits. The 
schedule usually is built into the fuel control by means of a three-dimensional 
cam. Most of the variables sensed by the fuel control affect the movement of 
the cam in one way or another. Due to the difficulties in installing and main- 
taining thermocouples in the area of the turbine, this empirical method of 
computing and controlling fuel flow is generally preferred, particularly in the 
larger gas turbine engines. 

The direct result of varying fuel flow is higher or lower combustion tempera- 
tures. If fuel flow is increased, the temperature of the air passing through the 
burners is increased, accompanied by a small increase in pressure that reacts 
on both the turbine and the compressor. The turbine receives added heat and 





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pressure energy. In the compressor, the effect is a small back pressure. The 
compressor rotors accelerate in spite of the added pressure which it must 
overcome, provided the pressure is not too great. The increased compressor 
speed forces additional air through the engine with the final result that thrust 
is increased. Reducing fuel flow starts a series of events, beginning with a 
lowering of combustion temperature and ending with less thrust. Fuel flow 
needs, which vary with changing compressor inlet conditions, are not left to 
the pilot’s judgment. To maintain the thrust, called for by the pilot, at a con- 
stant percentage of maximum thrust, the control automatically decreases fuel 
flow with decreasing compressor inlet temperature or pressure as altitude is 
gained. The reverse takes place during descent. 

The handling of the engine during steady state conditions is only a part of the 
responsibility of the fuel control. When the engine is accelerated, energy must 
be supplied to the turbine in excess of that necessary to maintain a constant 
rpm. However, if the fuel flow is increased too rapidly, an overrich mixture 
may be produced that will cause excessive turbine inlet temperatures or a rich 
blowout. Conversely, reducing the fuel flow too quickly during deceleration 
may result in a lean die-out because it is possible to reduce the fuel flow at a 
faster rate than the rate at which the compressor will reduce airflow to the 
burners. The fuel control must maintain engine operation within limits of 
fuel/air ratio, which will preclude the possibility of a flameout during either 
acceleration or deceleration. 

Another responsibility of the fuel control is the prevention of compressor 
stall. Stall is a characteristic of all airfoil sections, including those used in gas 
turbine engines. Stall is not a reflection of the engine’s reliability. Stall is a 
condition of unstable airflow through the compressor to which the engine is 
more susceptible at low compressor inlet temperatures. It is necessary to limit 
acceleration fuel flow schedules in order to avoid stall. At certain compressor 
speeds and inlet air temperatures, care must be taken not to introduce fuel 
into the burner section in such a manner that a critically high pressure will be 
produced before rpm, airflow and air pressure increase sufficiently to take care 
of the situation. Under such conditions, the flow of air through the compressor 
slows down and the compressor blades stall. The fuel flow schedule of the fuel 
control is designed to guard against this condition. 

: Although the basic requirements which a fuel control must satisfy apply, 

in general, to all gas turbine engines, the means by which individual controls 
satisfy these needs cannot be conveniently generalized. There are as many 
variations in controls as there are fuel control manufacturers. Each develop- 
ment group analyzes its own problems and pursues a course based on its best 
judgment and experience. Also, individual differences occur because of the 
operating features of specific engine types or models. 

Fuel controls can be divided into two basic groups, hydromechanical and 
electronic. In reality, the latter is a combination of the two. Although each 
type of fuel control has its particular advantages, most controls in use today 
are the completely hydromechanical type. Regardless of type, all controls 
accomplish essentially the same functions, although some sense more engine 
variables than others. 




PWA OPER. |[NSTR. 200 

At best, fuel controls are extremely complicated devices. The hydromechan- 
ical types are composed of speed governors, servo systems, sleeve and pilot 
valves, feed-back or follow-up devices, metering systems, and the like. Elec- 
tronic fuel controls are a maze of thermocouples, amplifiers, relays, electrical 
servo systems, switches and solenoids. Each individual fuel control must be 
carefully studied to be understood. However, the average pilot, mechanic or 
flight crew member ordinarily does not have a need for such complete knowl- 
edge. For those who must pursue the subject of fuel controls more thoroughly, 
excellent descriptive manuals, cutaway drawings and sketches for each of the 






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various controls are available. These are usually published by the engine 
manufacturer. Figure 2-41, showing the Hamilton Standard J FC 12-11 fuel 
control, is representative of cutaway drawings of this type. It will serve to 
illustrate the complexity of fuel control operation and is included here partic- 
ularly for the benefit of those who may wish to trace the internal mechanisms 
of a typical fuel control. } 

Figure 2-42 schematically illustrates a very simplified version of a typical 
fuel control for a turbojet engine. The sketch shows the basic fuel control 

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No Acceleration : 
GSanrner T 
= T+2 Constant 
To No Sense 
os = 

Trimming Adjustment 

(Changes Fulcrum pee 

Point of Lever) JOH : 

: . « ine. Ws ! 
: For any given 
We altitude & Tro 
To Tho or Pio Sense 1 oe 
& Compression 
“ oo 
5 te 
Ty2 or Pro Bellows and Cam 
) | (| Ta Cold Hot : 
5 lj I} ——~- jj —— - YOUR ee 
eae! Low High "N Power Lever Constont 

Tk2-Compressor inlet total temperature 

P+2-Compressor inlet total pressure 

N2-High pressure compressor RPM T2 oF Pt2 —» 

Pb-internal engine burner pressure Increase ——e TREND CURVES 

f-Fuel flow pounds per. hour (All curves are for steady state conditions 
PM except as noted) 

3% Not part of basic fuel control 


Figure 2—42. Basic Fuel Control Functions 


. JNSTR. 200 




functions but does not include such items as servomechanisms, feed-backs 
and solenoids. A study of the diagram will assist the reader in understanding 
what a fuel control does, rather than precisely how it operates. The curves 
included in Figure 2-42 illustrate the trends of the various engine variables 
which the fuel control endeavors to establish. By tracing the action of the 
push-pull rods on the sketch and the various cams which govern their effec- 
tive length, it is possible to follow the simulated action of the fuel control from 
the time that the fuel control receives a signal from the cockpit throttle. With 
a little imagination, it is possible to understand how varying compressor inlet 
pressure and/or temperature, rpm and burner pressure affect the amount of 
fuel metered to the engine for various flight conditions. The sketch portrays 
all of the effects in two dimensions only. In an actual fuel control, three- 
dimensional cams and other complicated devices take the place of the simple 
mechanisms shown here. 

Turboprop Fuel Controls and Propeller Governors 

A turboprop fuel control is similar to that of a turbojet engine in that it is 
usually hydromechanical and controls fuel flow with respect to compressor 
speed, compressor inlet temperature and engine internal pressure. The control 
operates according to a prescribed schedule which maintains a constant tur- 
bine inlet temperature consistent with the amount of power called for by the 
throttle setting. The control monitors burner pressure and decreases fuel flow 
as the ambient temperature and burner pressure decrease with increased alti- 
tude. Turboprop fuel controls differ from turbojet fuel controls by virtue of 
the fact that a propeller governor operates in conjunction with the turboprop 
fuel control. 

The turbojet engine produces thrust directly, since there is a specific relation- 
ship between compressor rpm and performance. The turboprop engine pro- 
duces thrust indirectly, since the compressor-turbine assembly furnishes torque 
to a propeller which, in turn, produces the major portion of the propulsive 
force which drives the aircraft. The turboprop fuel control and the propeller 
governor are connected and operate in coordination with one another. The 
throttle directs a signal from the cockpit to the fuel control for a specific 
amount of power from the engine. The fuel control and the propeller 
governor, together, establish the correct combination of rpm, fuel flow and 
propeller blade angle to create sufficient propeller thrust to provide the 
desired power. 

The operating schedule of the fuel control and propeller governor is usually 
divided into two regimes: one for flight and one for ground handling. For 
all airborne operation, the propeller blade angle and fuel flow for any given 
throttle setting are governed automatically according to a predetermined 
schedule. Below the Flight Idle throttle position, the coordinated rpm-blade 
angle schedule becomes incapable of handling the engine efficiently. Here, 
the ground handling, or Beta, range is encountered. In this region of the 
throttle quadrant, the propeller blade angle is not governed by the propeller 
governor but is controlled by the throttle position. A Ground Idle throttle 


_ ZT GT Ge 

PWA OPER. | NSTR. 200 

setting is provided for engine starting and operation on the ground. When 
the throttle is moved below the Ground Idle position, the pitch of the propel- 
ler blade becomes reversed to provide reverse thrust for rapid deceleration 
of the aircraft after landing. Blade angle pitch stops are usually provided in 
the propeller in order to maintain a minimum blade angle should a propeller 
malfunction occur under critical in-flight or take-off conditions. 

The varying amounts of torque required by changing flight conditions are 
absorbed through the expedient of increasing or decreasing the propeller 
blade angle. Compressor rpm and torque produce shaft horsepower which, in 
turn, varies with fuel flow which is established by the signal from the throttle 
to the fuel control and the propeller governor. The design engineer who deter- 
mines the operating schedule must consider the best combination of engine 
speed, specific fuel consumption and propeller performance. He must also take 
permissible temperature, pressure and rpm, as they relate to engine durability, 
into consideration. Both Ground Idle and Flight Idle power and rpm values 
may vary greatly with different types of fuel controls and different engine 
installations. These values are preset in the fuel control and propeller gover- 
nor schedule to produce predetermined performance characteristics for a 
particular airplane. Compressor stall during acceleration and deceleration is 
handled in much the same manner for the turboprop engine as that for a 
turbojet engine. When the engine is operating in the ground handling or 
propeller nongoverning (Beta) range, stall is usually prevented by overboard 
interstage compressor airbleeds. In the flight operating range, compressor 
stall prevention is a function of the fuel control and propeller governor sched- 
ule which is designed to avoid conditions that might tend to produce stall. 

Lubrication Systems 

Oil in a gas turbine engine serves the twofold purpose of cooling and lubricat- 
ing the bearings. A pressure oil system carries the oil directly to the points 
where it is needed. These are the compressor and turbine bearings and the 
accessory drive shaft bearings in a turbojet engine, and, additionally, the 
propeller shaft bearings, the reduction gearing and the torquemeter in a 
turboprop engine. The lubrication method most generally used is known as 
a “calibrated” system because each bearing has its oil specifically controlled 
by a calibrated orifice which provides the proper oil flow at all engine 
operating speeds. 

Figure 2-43 represents a typical lubrication system for a dual axial compressor 
engine. The lubrication system for a centrifugal compressor engine is some- 
what similar. Oil is taken from the oil storage tank by a gear-type, pressure 
pump and is passed through a screen at the pump outlet to oil lines which 
carry it to various points in the engine. Return oil is scavenged either at each 
bearing or at collecting points by gear-type, scavenge pumps, after which it 
passes through a suitable cooler to the storage tank. Oil coolers may be either 
the air-oil type or, on some fighter aircraft, a combination of an air-oil cooler 
and a fuel-oil cooler. 






manufacturer) By 4 





Figure 2—43. Typical Lubrication System For a Dual Axial Compressor Engine 

As it re-enters the storage tank, the oil is passed through a deaerator to sep- 
arate most of the air from the scavenged oil. Can-type deaerators are most 
commonly used and are usually an integral part of the storage tank. The air 
released is vented overboard. The oil storage tank must contain baffling to 
prevent re-aeration of the oil. The 
lubrication system must be provided 
with a breather and pressurizing 
valve at the oil tank (Figure 2-44) 
to keep the oil under pressure at 
high altitude. This consists of a 
spring and bellows operated poppet 
valve and a spring-loaded blowoff 
valve which maintain the pressure 
at some predetermined value. At sea 
Figure 2—44. Typical Lubrication level pressure, the bellows operated 
System Breather and Pressurizing Valve valve is open. It closes gradually, 
through operation of the bellows, 
and a sufficiently high pressure is maintained in the oil system to assure proper 
operation of the system as altitude is gained. The spring-loaded blowoff valve 
acts as a pressure relief valve for the entire system and will open only if pres- 
sure above a predetermined maximum builds up within the system. 

Because only the compressor and turbine bearings and the accessory drive 
shaft bearings require lubrication, turbojet engine oil consumption is rel- 
atively low as compared with that of a reciprocating engine. Slightly less than 
half a gallon per hour is about normal for engines of intermediate size. There 
is more to lubricate in a turboprop engine, so the consumption is higher, size 



| on | 




SS Ga 



r for size, than for a turbojet. Most gas turbines do not require preheating 
except in the very coldest of weather and, because of the characteristics of 
their oil system, are ready for full thrust or power operation almost imme- 
diately after the engine is started. In a turboprop engine, the high loading on 
the reduction gearing makes the lubrication of this area rather critical. Pre- 
oiling may therefore be necessary if the engine has not been operated for an 
appreciable length of time. This is accomplished by accelerating the engine 
with the starter, without ignition or fuel, then allowing the compressor to 
decelerate before commencing the normal starting sequence. 

Ignition Systems 

Although all combustion engines can be ignited quite easily under ideal con- 
ditions, gas turbines normally operate at high altitudes where conditions for 
an in-flight engine relight in the event of a flameout are far from ideal. The 
low temperatures encountered at high altitude cause a decrease in fuel volatil- 
ity which makes it difficult to ignite the fuel charge. It is necessary to have not 
only a very high voltage to jump a wide igniter plug spark gap but also a 
spark of high heat intensity. The high-energy, capacitor-type ignition system 
has been more or less universally accepted for gas turbine engines because it 
provides both high voltage and an exceptionally hot spark which covers a 
large area. Excellent chances of igniting the fuel/air mixture are assured at 
reasonably high altitudes. 

— = & = © 

The ignition system described here is considered typical. Others may vary 
slightly, especially in their capacity, but operate in essentially the same man- 
ner. Ignition systems are rated by capacity and may be, for instance, one-, 
two- or four-joule units. One joule is the unit of work or energy eXpended in 
one second by an electric current of one ampere in a resistance of one ohm. 
That is, one joule equals one watt per second. The typical ignition system is a 
four-joule unit, operated by a standard 24-volt (16-30 volt range) aircraft 
electrical system. 

The combustion chambers of can-type and can-annular-type gas turbine en- 
gines are interconnected by flame tubes in such a manner that a flame started 
in one chamber will spread rapidly to the others. Annular-type chambers 
need no such arrangement. Normally, gas turbine engines are provided with 
two igniter plugs, sometimes called spark igniters. In the case of can-type and 
can-annular-type burner chambers, the plugs are located in separate cham- 
bers. Igniter plugs serve a purpose similar to the spark plugs in a reciprocat- 
ing engine. However, unlike a reciprocating engine, operation of the ignition 
system and the igniter plugs is necessary only for a short period during the 
engine starting cycle. The typical ignition system consists of two identical, 
independent ignition units with a common electrical power source, the air- 
craft battery. To provide a safety factor, there is one ignition unit for each 
of the two igniter plugs. Since the circuits to each of the plugs are identical, 
only one will be considered in detail (Figure 2-45). 

ZZ TZ | 

aA The first part of the ignition system is an input filter to eliminate radio inter- 
ference. This is required since the aircraft battery is used as a source of power 


PWA OPER. | NSTR. 200 


for both the aircraft radio equipment and the ignition. Radio noise originates 
basically with alternating current. The input filter permits the flow of direct 
current to the ignition in one direction and weakens alternating or pulsating 
current in the opposite direction toward the aircraft radio. 

Together with the input filter, the primary part of the ignition system is known 
as the ignition exciter unit. When the aircraft ignition switch is turned on, 
direct current is supplied through the filter to an electric motor which operates 
two cams. One is a multilobe cam in the exciter unit. The other is a single-lobe 
cam for the ignition compositor circuit. Both cams open and close breaker 
points to supply intermittent current to transformers. Referring to Figure 
2-45, the multilobe cam delivers a pulsating current to a low-tension trans- 

Igniter Plug 


High Altitude Terminals 

Ignition 2 
-G)—Compatiter —__ 
| High-Tension | 
| Transformer | 
| | | 
: | 
Ignition Exciter 
OES ET ee ee Lee Se ee MRRP EN 0M vi 
| Condenser : 
| T | 
= | 
| Resistor | 
| Single-Lobe Cam : 
| -_ | 
| Large Storage | 
| input Filter T —— Selenium 
| Bese Psy = Rectifier 
| | 
ir 7 
| | 
| | | 
| | Multi-Lobe Cam Low ~ Tiasion 
| | Transformer 
ha | 
i | 
PL Condenser | 


Figure 2—45. Typical High-energy, Capacitor-type Ignition System 

former. A condenser prevents arcing across the breaker points. The trans- 
former increases the 24-volt input voltage to about 2,000 volts in the system. 
This voltage is then passed through a selenium rectifier which acts as a One- 



way check valve, allowing the flow of current into a storage condenser but 
preventing any flow in return. The storage condenser stores up a huge amount 
of current each time that the breaker points open. At the same time that the 
storage condenser is being charged, the single-lobe cam in the compositor 
closes its breaker point, permitting one pulse of direct current to flow through 
the primary coil of a high-tension transformer. The high-tension transformer 
increases the voltage to about 28,000 volts. This very high voltage causes a 
“trigger spark” to jump the wide spark gap in the igniter plug. Once the path 
across the igniter plug is bridged, a path of lower resistance is established for 
the discharge of the greater amount of electrical energy stored in the storage 
condenser. The relatively low voltage in the storage condenser, although 
capable of producing a very hot spark, is not sufficient, in itself, to bridge the 
igniter gap until a path is provided by the trigger or leader spark. The com- 
bined result is a very intense spot of heat which is capable of quickly igniting 
the fuel/air mixture in the burner chamber. 

To avoid the danger of an internal explosion during engine starting, good 
practice dictates that the engine should first be turned over by the starter and 
permitted to attain sufficient speed to purge itself of any combustibles which 
may be present before the ignition is turned on. The ignition is then turned 
on before fuel is sprayed into the engine. The ignition cycle takes place several 
times per second and continues to operate as long as the ignition switch is on. 
It is common practice to install a timing mechanism in the starting circuit, 
which will limit the time that the ignition is energized. An alternate means of 
accomplishing this is to incorporate a cutoff switch in the throttle quadrant 
which will turn the ignition off when the throttle is advanced above the start- 
ing or Idle position. Ignition units are usually contained in hermetically sealed 
boxes. In the event of a malfunction, it is necessary to replace the entire unit. 
High-altitude shielded cable is employed throughout the system. 

Igniter Plugs 

The electrode of gas turbine engine igniter plugs must be able to accommo- 
date a current of much higher energy than the electrode of conventional spark 
plugs is capable of handling. Although the high-energy current causes more 
rapid igniter electrode erosion than that encountered in reciprocating engine 
spark plugs, it is not of any consequence because of the relatively short time 
that a turbine engine ignition system is in operation. It does, however, consti- 
tute the reason for not operating gas turbine ignition systems any longer than 
is absolutely necessary. Igniter plug gaps are large in comparison with those 
of conventional spark plugs because the operating pressure at which the plug 
is fired is much lower than that of a reciprocating engine. The fouling common 
to reciprocating engine spark plugs is held to a minimum in igniter plugs be- 
cause of their high-intensity spark. 

Most igniter plugs are of the annular-gap type (Figure 2-46), although con- 
strained-gap plugs are used (Figure 2-47) in some engines. Normally, the 





annular-gap plug projects slightly into the combustion chamber liner in 
order to provide an effective spark. This is sometimes called a “long reach” ig- 
niter. The spark of the constrained- 

gap plug does not closely follow ities 
the face of the plug; instead, it ——— ee 
tends to jump in an arc which car- 

ries it beyond the face of the cham- Figure 2—46. 
ber liner. The constrained-gap plug Porcelain 
need not project into the liner, with rei 
the result that the electrode oper- 
ates at a cooler temperature than 
that of the annular-gap plug. 

of ; 
— FS 

Figure 2—47. 
Engine Cooling 

The intense heat generated when fuel and air are burned necessitates that 
some means of cooling be provided for all internal combustion engines. 
Reciprocating engines are cooled either by passing air over fins attached to 
the cylinders or by passing a liquid coolant through jackets which surround 
the cylinders. The problem is made easier by the fact that combustion takes 
place only during every fourth stroke of a four-cycle engine. In the case of 
a gas turbine, the burning process is continuous, and nearly all of the cool- 
ing air must be passed through the inside of the engine. If only enough air 
were admitted to the engine to provide an ideal air/fuel ratio of 15:1, internal 
temperatures would skyrocket to more than 4,000°F. In practice, a large 
amount of air, in excess of the ideal ratio, is admitted to the engine. The large 
surplus of air serves to cool the hot sections of the engine to acceptable temp- 
eratures ranging from 1,100° to 1,500°F. Figure 2-48 illustrates the approx- 
imate engine outer case (skin) temperatures encountered in a properly cooled 
dual axial compressor turbojet engine. Due to the effect of cooling, the temp- 
eratures of the outside of the case are considerably less than those encountered 
within the engine. The hottest spot occurs opposite the entrance to the first 
Stage of the turbine. Although the gases have begun to cool a little at this 
point, the conductivity of the metal in the case carries the heat directly to the 
outside skin. 

The air passing through the engine serves to cool the combustion chamber 
burner cans or liners. The cans are constructed in a manner which serves to 
induce a thin, fast-moving film of air over both the inner and outer surface 
of the can or liner. Can-annular-type burners frequently are provided with a 
center tube to lead cooling air into the center of the burner to promote high 
combustion efficiency and rapid dilution of the hot combustion gases while 
minimizing pressure losses. The Pratt & Whitney Aircraft J57 is an example 
of an engine with a burner of this type. In all types of gas turbines, large 
amounts of relatively cool air join and mix with the burned gases aft of the 
burners to cool the hot gases just before they enter the turbines. 



OF ns Cooling air inlets are frequently provided around the exterior of the engine to 
permit the entrance of air to cool the turbine case, the bearings and the tur- 
bine nozzle. In some instances, internal air is bled from the engine compressor 

f section and is vented to the bearings and other parts of the engine. Air vented 
into or from the engine is ejected into the exhaust stream. When an accessory 

r case is mounted at the front of the engine, it is cooled by inlet air. When lo- 
cated on the side of the engine, the case is cooled by a flow of outside air 
ducted through it. 

The exterior of the engine and the engine nacelle are cooled by passing air 
P between the case and the shell of the nacelle (Figure 2-49). The engine com- 
partment is frequently divided into two sections. The forward section is built 
around the engine air inlet duct; the aft section is built around the engine, it- 
self. A fumeproof seal is provided between the two sections. The advantage of 
such an arrangement is that fumes from possible leaks in the fuel and oil lines 
P contained in the forward section cannot become ignited by contact with the 
| hot sections of the engine, itself. In flight, ram air provides ample cooling of 
the two compartments. On the ground, air circulation is provided by the 
effect of reduced pressure at the rear of the engine compartment, produced by 
gases flowing from the exhaust nozzle. 


P ° 
r 5 
F £ 
P 2 ven eee eee 
; oo Se 
S) | 
- Fas > th 
P e|F Exhaust 
r Low Pressure | Burner ®)@/. IDuct & 
Compressor 55 5|/5\/o 
DD 2 | m| >|Nozzle 
o 9 2| ¢ 
F a2 a \a}8 
£ im So 
mo Oo =| 
=? 25's 
rm Figure 2—48. Typical Outer Case Temperatures For a Dual Axial Compressor 
| Turbojet Engine 




Fume Proof Seal 

— Aft Compartment 

*lelitatems D A helaaiele men 


Inlet Duct 

ee) Forward Compartment 

| Tire ae , 

Starter | Case Cooling Air. 



Figure 2—49. Typical Engine Nacelle Cooling Arrangement 

Engine Insulation Blankets 

To reduce the temperature of the structure in the vicinity of the exhaust duct 
or afterburner, and to eliminate the possibility of gasoline or oil coming in 
contact with the hot parts of the engine, it is sometimes necessary to provide 
insulation on the exhaust duct of gas turbine engines. As shown in Figure 
2-48, the exhaust duct surface temperature runs quite high. Much higher tem- 
peratures might be obtained during a hot engine start. Metal surfaces at this 
temperature transmit a significant amount of heat by radiation as well as by 
conduction and convection. Because of this, it has been found desirable to 
incorporate radiation shields of aluminum foil as well as blankets of low- 
conductance material in some engines. A typical insulation blanket and the 
temperatures obtained at various locations are shown in Figure 2-50. This 
blanket contains Fiberglas as the low-conductance material and aluminum 
foil as the radiation shield. The blanket is suitably covered so that it does not 
become oil-soaked. 

Water or Coolant Injection 

The sensitivity of gas turbine engines to compressor inlet temperature results 
in appreciable loss of the thrust (or power, in the case of a turboprop) which 
is available for take-off on a hot day. It is frequently necessary, therefore, to 
provide some means of thrust augmentation for nonafterburning engines dur- 
ing take-off in hot weather. Ten to fifteen per cent additional thrust (power) 
can be gained by injecting water, or a mixture of water and alcohol, into the 
engine, either at the compressor air inlet or at some other point, such as the 
diffuser case. In a reciprocating engine, during power augmentation by means 
of water injection, the water acts primarily as a detonation suppressor and a 
cylinder-charge coolant. Induction air cooling is secondary. Higher take-off 
horsepower results chiefly because when water, or a mixture of water and al- 
cohol, is added, the engine can operate at the fuel/air ratio that will produce 
“best power.” Sometimes a higher manifold pressure may be obtained than 
would otherwise be possible without experiencing detonation. Gas turbines, 


Biren | | 
lf > Engine 
low Pressure gh : 
et | pi — 
. | 





PWA OPER. | NSTR. 200 

Outer Engine Compartment Casing 
Stainless Steel Shroud-350° F 

Pee eS 

Cooling Air ——® I20 F 

o7e767670'6 0.0.0 2.9% % ©9999, 9.%5terar ere #0 0.0.0, 0,0 9,050 steer ere eo. 
Saar te ee ar ad Oe ae MO ddd 

Aluminum Foil 

Silver Foil 

Exhaust Duct 900F 

CO i aad te Oe ie Oe MM 

Exhaust Gas —> |OOOF 

Figure 2—50. Typical Engine Insulation Blanket 

however, have. no detonation difficulties. When a liquid coolant is added, 
thrust or power augmentation is obtained principally by cooling the air enter- 
ing the engine by means of vaporization of the liquid introduced into the air- 
stream. Cooling the air has the effect of reducing the compressor inlet tem- 
perature. The reduction in temperature increases the air density and the mass 
airflow. More and cooler air to the burners permits more fuel to be burned 
before limiting turbine inlet temperatures are reached, which, in turn, means 
more thrust. 

The effect upon the engine’s thrust depends upon the type of coolant used, the 
proportion of the ingredients and the quantity of the coolant flow. In order to 
obtain effective cooling, a liquid with a high heat of evaporation is required. 
About the only suitable liquids available are methyl, or ethyl, alcohol and 
water. Because water has the highest heat of vaporization, more cooling can 
be obtained with water than with an equal mass of either of the other two. In 
addition, alcohol is expensive. Consequently, water, by itself, is the most desir- 
able coolant. Alcohol is added occasionally in varying proportions, either to 
lower the freezing point of the coolant or to eliminate the need for separate 
enrichment of the fuel mixture, which might be necessary if only pure water 
were used. When alcohol is added, some small amount of additional thrust 
may be produced as the alcohol is burned. However, the efficiency of the com- 
bustion of the alcohol is usually quite low. The heating value of methyl or 
ethyl alcohol is only about half that of kerosene or gasoline. Most of the flow 
of the alcohol/air mixture will not pass through that part of the combustion 
zone where temperatures are high enough to support efficient combustion of 
the weak alcohol/air mixture. 

When water is used as a coolant, it should be nearly pure chemically, prefer- 
ably containing not more than 0.005% total solids. Dissolved mineral content 
in the water causes a rapid buildup of deposits on the compressor blades and 
vanes. Although deposits have only a small influence on engine operation at 
the “wet” (water injection) take-off condition, rapid deterioration of the 
engine “dry” (without water injection) thrust occurs which is noticeable at 
any given engine speed. This results in an increase in the frequency of re- 




quired cleaning of the compressor, either in the field or by overhaul. In the 
case of dual axial compressor engines, more frequent engine trimming be- 

comes necessary. The deposits are most noticeable on the high pressure com- 

pressor. It is highly advisable that water demineralizing equipment be used 
for servicing aircraft water injection systems. 

Coolant injection is usually accomplished by spraying the liquid into the en- 
gine at the compressor air inlet. The typical injection system (Figures 2-5] and 



= 7 itt 
oe tema | PREes- || _ 
y = j SURIZING j TO Py2 — — — 
WATER & oumP 
! re | TT tae datiacdebates g x IGNITION SWITCH 


OIL TEMPERATURE ©- ait ide shales 

O14. pressune ©-- sia 

as kl 

[a | 

Figure 2—51. Typical System for Water Injection at the Compressor Inlet 


he ste Te tt ee es sl am 

eee ae a ee ee 

11Q |uear (Qs |b Heat 
7 i 
ee es 
“Wess SSSSee eee 




Wy pees errerr 

a B, 



| 2 







J | | 
S 2 = 



| ” 


Gu SHUT-OFF VALVE (on ruee sysTeM 
cuosen] SHUT © 







and the Diffuser Case 



—_™ mm Ee Dae ee weep ee = 




2-52) contains a coolant regulator, a shutoff valve, a filter and a coolant in- 
jection pump located at the aircraft coolant tank. The pump is operated either 
by high pressure air, bled from the engine, or by a separate electric motor. The 
water shutoff valve is connected to the fuel control electrically so that the in- 
jection system will not be activated until the throttle has first been placed in 
the take-off position. The aircraft coolant tank must be provided with a means 
for draining, to prevent freezing at high altitude in the event that all of the 

Water Press. Gage 

oo Pa eee SF 
| |QDiffuser Water Z-Water In 
“  Finjection RegulatorsJee & Screen 


Diffuser Water 
Manifold And 

Water Injection 
Reset Connection 

Compressor Inlet 
Water Injection 

7? «Compressor 
Inlet Spray 

Figure 2—52. Compressor Inlet and Diffuser Case 
Water Injection Regulators 

coolant is not used during take-off. In some instances, the injection may be 
accomplished at the engine diffuser case. Although less efficient, diffuser case 
injection has the advantage of requiring less change in the basic engine design 
when coolant injection is added, and avoids the possible problem of deposits 
forming on the compressor blades. To be effective, water or coolant should 
be in excess of a coolant/air ratio of 1:100, by weight. The limiting amount 
of water or coolant which may be used is roughly in the ratio of 5:100 for an 
engine compression ratio of 5:1. Beyond this, complete evaporation in the 
compressor becomes either difficult or impossible. 

Protection Against Icing 

Although gas turbine engines have no carburetors and no attendant car- 
buretor-icing problems, they are not immune to the effect of freezing mois- 
ture. Under icing conditions, axial compressor engines are seriously affected 
by the formation of ice on the compressor inlet guide vanes. All turbine en- 
gines equipped with nonretractable air inlet screens are very susceptible to 
icing. Ice forms on the guide vanes or inlet screen and restricts the flow of 
inlet air. This is indicated by a loss of thrust and a rapid rise in exhaust gas 
temperature. As the airflow decreases, the fuel /air ratio increases, which, in 
turn, raises the turbine inlet and discharge temperatures. The fuel control at- 
tempts to correct any loss in engine rpm by adding more fuel, which aggravates 
the condition. For a given icing condition, small engines with close spacing 
of the inlet guide vanes are more seriously affected than large engines. 

The inlet guide vanes can be heated to prevent the formation of ice. However, 
the only way to eliminate the formation of ice on an air inlet screen which 





cannot be retracted is to avoid flying into such conditions. Engine operation 
under icing conditions is a subject in itself which will be discussed later. Cen- 
trifugal compressor engines, whether equipped with retractable screens or 
having no screens at all, are relatively free from the danger of ice collecting 
at the compressor inlet. The formation of airplane wing ice is much more 
serious than engine icing and is the limiting factor governing flight of air- 
craft equipped with this type of engine. 

The inlet guide vanes and the inlet struts of axial compressor engines are 
usually hollow. Hot, high-pressure air is bled from the rear of the engine com- 
pressor and is ducted through an anti-icing system control valve to the hollow 
sections of the inlet struts and guide vanes. The heat provided prevents the 
adhesion of ice. Because such a system may not melt ice, once it is formed, 
icing conditions should be anticipated in advance. An anti-icing regulator is 
usually incorporated which automatically regulates the flow of anti-icing air 
with changing compressor inlet temperature. Anti-icing systems cause some 
reduction in thrust and are used only when needed. The control valve nor- 
mally operates a warning light on the instrument panel, which indicates that 
the anti-icing system is in operation. 

Turbine engine fuel systems in which a low-pressure filter is incorporated may 
experience icing of the filter element if the fuel in the aircraft fuel tanks be- 
comes severely chilled. Such systems are often equipped with a small, alcohol 
tank. A tap is provided to permit injection of the alcohol (or similar deicing 
fluid) into the fuel line to assist in removal of ice from the low-pressure filter 
screen. An alternate method is to supply heat to the fuel, ahead of the filter. 
The aircraft low pressure fuel gage is ordinarily connected to a pressure 
switch which operates a warning light to give the pilot visual indication that 
the fuel pressure is below the allowable minimum, indicating that ice is 
forming at the filter. 

Another method of deicing parts of the fuel control, such as a servo line, 
which, in some engines, may be subject to ice formation, is by use of hot com- 
pressor bleed air. A solenoid-operated, heat exchanger valve is provided in the 
engine to supply hot bleed air to warm the ice-susceptible engine part, when- 
ever necessary. 

Turboprop Asymmetric Thrust Control Devices 

Studies have shown that in cases of simple loss of power, such as might result 
from lack of fuel, the turboprop airplane is not appreciably different from 
the reciprocating engine-powered airplane. Regardless of engine type, how- 
ever, the problem of asymmetric thrust from an inoperative powerplant be- 
comes aggravated with increased engine power, especially at the high air- 
speeds at which a turboprop aircraft operates. For cases involving possible 
multiple failure of turboprop engines or propellers, some means of automat- 
ically limiting the drag resulting from an inoperative engine is highly desirable 
to assure flight safety. Flat propeller blade angles at low pitch stops can pro- 
duce high rotational speeds in a windmilling turboprop engine which, in turn, 
induces high asymmetric thrust. 



[=> TT 

_ eee ae = [6S 

a = we 



Unbalance due to thrust loss and windmilling drag Asymmetric thrust (Figure 2-53) may be de- 
fined as the unbalanced thrust between 
either side of a multiengine airplane, re- 
sulting from the combination of loss of pro- 
pulsive thrust and the windmilling propeller 
drag caused by an inoperative powerplant. 
— It will be greatest in the case of a failed out- 
—" board engine. Windmilling drag is caused 
primarily by low blade angles and can be 
largely eliminated if the propeller pitch is 
een cee” ~~ not allowed to go below the take-off blade 
er angle. In the case of free turbine and dual 
axial compressor engines, it is possible to 
Figure 2—53. Asymmetric Thrust install low pitch stops set at the take-off 
blade angle, such as are used on recipro- 
cating engines to prevent high windmilling drag. With the single-compressor, 
directly-coupled engine, however, the almost constant speed of operation 
requires a propeller blade angle or pitch that is usually 10 to 15 degrees 
lower than the take-off blade angle, in order to provide zero thrust at landing 
speeds. Although low pitch stops can, in part, be relied upon to suppress 
windmilling drag, automatic means should be provided to prevent the blades 
from going to a flatter pitch and to start them moving toward the feathered 
position in the event of an engine failure. Such a device eliminates the neces- 
sity of relying entirely upon the alertness and quick action of the pilot to 
start the feathering operation during an emergency. A device of this kind 
assures that the propeller of the malfunctioning engine is the propeller 
that is feathered. 


To be completely satisfactory, an automatic antiwindmilling drag system 
must be absolutely reliable, must be capable of being checked before flight, 
and must be designed to ensure that the propeller blades will return to their 
normal operating position if power is re-established following a momentary 
engine power stoppage. The engine must not be committed to a shutdown 
should gusty flight conditions or too rapid a throttle movement cause a brief 
loss of power. 

Automatic decoupling systems and the addition of low pitch propeller stops 
which can be released in flight have been employed to solve the windmilling 
drag problem in directly coupled turboprop engines. However, such systems 
have excessive complexity. A more effective and reliable method of suppress- 
ing the windmilling drag from an inoperative powerplant, prior to feathering 
the propeller, has been developed which is known as the negative torque 
control (NTC) system (Figure 2-54). When a large power reduction or mal- 
function occurs and a windmilling propeller starts to drive the engine, a 
negative torque condition is set up. As the direction of torque in the propeller 
shaft reverses, the reduction ring gear (which is usually either the torque- 
meter ring gear or a separate NTC ring gear) in the engine reduction gear 
housing reverses axial direction and moves forward in its splines. The forward 
motion of the ring gear provides the force which operates the NTC. As the 
reduction ring gear changes position, it moves a small plunger which, in 




turn, actuates either a hydraulic valve or an electric switch, thus producing 
a signal to the propeller mechanism. The signal starts the blades turning to- 
ward their feathered posi- 
tion. The system can be a- 
dapted to either electrical 
or hydraulic propellers. 

The negative torque con- 

‘ trol is frequently not an ———— 
os automatic feathering de- —— 
; * __ #ice. In some installations, | ——— 
re the pilot must feather the Governing Propeller Ponction 1 
| engine manually as soon _pgcrease ; oe a Reduction 

eo oe 
as it has been positively — Piteh , LR 

determined that an engine 
failure has occurred. If — fropeller =“ 
power Is restored win. —_ 
time prior to manual Pitch 
feathering, the NTC sig- roe: testes 
nal ceases and control of 

the propeller pitch is au- ELECTING. SYSTEM 
tomatically returned to the To increase pitch mechanism only To governing propeller function 
propeller governing sys- 4| 

tem. Normal operation is Figure 2—54. 
resumed without atten- NTC For Turbo- 
tion from the pilot. prop Engines 

Plunger moves 
forward with 
negative torque 


Turboprop Engine Brakes 

Because there is very little friction within the engine, especially at low engine 
speeds, and no piston compression to overcome, the propeller of a turboprop 
engine on a parked airplane has a tendency to windmill easily in the lightest 
breeze. Consequently, it is usually good practice to feather the propeller on 
very large engines whenever the engine is shut down on the ground. As a fur- 
ther precaution against windmilling, turboprop engines are usually fitted with 
some manner of parking brake for the engine. This is normally a small shoe- 
type of brake, fitted to the starter drive shaft or to one of the accessory drives. 
It is méchanically actuated from the cockpit and is primarily intended for 
ground parking use only. However, in some installations, the brake may be of 
sufficient size to permit limited use to prevent windmilling of the feathered 
propeller of an inoperative engine in flight. Under such circumstances, ex- 
treme caution is necessary, since the large amounts of propeller torque gen- 
erated at high airspeed may easily cause the brake to burn out. The recom- 
mendation in the Flight Handbook pertaining to the use of the brake should 
be strictly observed. Even at moderate airspeeds, a propeller windmilling so 
Slowly that the blades may be counted will absorb sufficient torque from the 
airstream to overload the brake. When in-flight use is permissible, the brake is 
usually designed to prevent the propeller from starting to windmill, not to stop 
it from windmilling once it is doing so. When the aircraft is on the ground, 
the brake should be applied at all times when the engine is not running. 




Section ITI 

Gas Twrhine Engine Porte an 

This section deals with the internal processes that take place within a gas 
turbine engine and the laws which govern them. Section II was primarily 
concerned with the description of the engine. Section III will explain as 
simply as possible its performance characteristics. 


More definitions and terms are needed to supplement those given in Section 
I in describing basic thermodynamic principles and some of the physical 
characteristics of gases. All of the following terms are important if one is to 
understand gas turbine engine performance. 

Heat and Temperature 

Gas turbine engines produce work in proportion to the amount of heat re- 
leased internally. For this reason, it is necessary to study the production 
of heat. within the engine, most of which is obtained by burning fuel, although 
some heat originates when air is compressed in the compressor. 

Normal ranges of temperature can be measured directly by simple thermo- 
meters or thermocouples. These indicate the level of molecular activity (in- 

ternal energy) of a body or substance, regardless of its total heat content. . 

Everyone is familiar with the ordinary thermometer which indicates the 
temperature of air or water. In contrast to temperature, heat cannot be 
measured directly but must be calculated from three known quantities. These 
are temperature, mass (or weight) and specific heat. 

Although the exact nature of heat is not known, arbitrary standards have 
been internationally agreed upon by which changes in heat content can be 
calculated accurately. The standard in the English System is the British 






Thermal Unit or BTU. In the Metric System, the gram-calorie is used. One 
BTU is defined as the amount of heat required to raise the temperature of 
one pound of pure water one degree from 59°F (15°C). The specific heat of 
water under these conditions is arbitrarily fixed at 1.0000. All other sub- 
stances are related to pure water through the values of specific heat found for 
each substance by experiment. A definition of heat, then, would be “that 
which causes a change in the temperature of a body.” The measure of heat 
generally used in the United States is the BTU. 

To illustrate further the difference be- Two Iron Blocks at the Same Temperature 
tween heat and temperature, imagine Both Blocks Have a Heat Content of IOBTU/LB 
two blocks of iron, one weighing 10 
pounds and the other 1000 pounds 
(Figure 3-1). Both are at the same tem- 
perature. The larger block, however, 

| le 
has 100 times the “heat” of the smaller cl 
because it has 100 times the mass. vy : 
Imagine, also, two turbojet engines, Weight 1000 Lb 
one using 10,000 pounds of fuel per Small Block Has 10x |O=Total BTU 
hour and the other using 1000 pounds Large Block Has 10x |OOO=Total BTU 
per hour. Both engines are operating 
at the same turbine inlet temperature. Figure 3—1. Heat Content 
However, the larger engine can do and Weight 

approximately 10 times the work of the smaller because 10 times more heat 
is released at the same temperature. 

From these examples, it can be seen that heat and temperature, although 
related, are not the same thing. The relationship between the four major 
‘ * ‘i a“ thermometer scales is shown in F igure 
~ Il 3-2. Two of them, Fahrenheit and 
Celsius (Centigrade), are the basic 
thermometers commonly used. By a 

10.0") | 222 Water to Steam 72°) | 212° comparatively recent decision of the 
International Civil Aviation Organiza- 

O° | #| 273° Ice Melts 492°| | 52°. tion (ICAO), the system of tempera- 
ture gradient commonly known as 
“Centigrade” in ‘the United States is 
henceforth to be internationally re- 
ferred to as Celsius, for aeronautical 
~agoe USC; after its originator. The other two 
) scales, Rankine and Kelvin, are some- 
times called “Fahrenheit — Absolute” 

-273°}8| O° Absolute Zero o° 

Celsius or Fahrenheit f : 
Centigrade and “Celsius (Centigrade )— Absolute” 
Figure 3—2. Comparison of scales, respectively. The latter two use 
Thermometer Scales the more common scales by simply 

adding a constant, which for Rankine 
is 460° and for Kelvin is 273°. The absolute scales are used in thermody- 
namic work, since they are better working tools than the common scales. 


—— ee 


—f Sie \ ag —_ =D = GD TD a | 


“= 5 y 

—— = 2S =D = = == == 






Mass and Specific Heat 

Mass is a basic property of matter. It is called “weight” when it is in a field 
of gravity like that of the earth. A block of iron weighing 100 pounds on 
the earth will weigh almost nothing out in interplanetary space because the 
attraction of gravity for the body which weighed 100 pounds on earth is 

practically zero. 

However, the force required to accelerate the block from a standing start 
to a given speed (assuming a frictionless track and no air resistance on the 
earth) will be the same at the earth’s surface as out in space. This force is 
needed to overcome the inertia and accelerate (or decelerate, as the case 
may be) the mass which, at sea level on the earth, is a 100-pound block of 
iron. Likewise, the number of BTU required to heat or cool this block from 
one temperature to another will be the same in space as on the earth, given 
the same conditions surrounding the block. Mass, then, is the amount of 
material in a body. 

When specific heat is used, it must be remembered that all substances are 
related to pure water at standard conditions, and that specific heat is the 
ratio of the amount of heat required to raise the temperature of a one-pound 
mass of a given substance 1°F to that required to raise one pound of water 
1°F (from 59°F). To illustrate the relation of other materials to water, sup- 
pose some substance has a specific heat of 0.500 at 75°F. This means that 
only half as much heat is required to raise its temperature to 76°F as is 
needed to raise the temperature of an equal mass of water from 59° to 60°F. 
The properties of mass and specific heat apply to all matter, whether solid, 
liquid or gaseous. 

Pressure Ratio 

Any machine that works with compressed or expanded gases works over a 
range of pressure ratios. The turbojet engine, for instance, takes air in the 
front, compresses the air to a given pressure and then permits the pressure 
to drop back to atmospheric as the gases are discharged from the engine at 
the exhaust nozzle. Each time the pressure is raised or lowered, a pressure 
ratio is involved. Thus, one 
speaks of the compression ratio 
of the gas turbine compressor, 
meaning the outlet pressure di- 
vided by the inlet pressure. In 
reciprocating engines, the pres- 
sure ratios are developed both 
by the supercharger and by the 
piston rising in the cylinder as it 
compresses the charge within, 
although this latter compression 
ratio is, strictly speaking, the | N 
result of a volume change. This Figure 3—3. Pressure Ratio ina 
is shown in Figure 3-3. Reciprocating Engine 

Ratio '!09/20=5:1 





Temperature Ratio 

A temperature ratio is identical to a pressure ratio except that it applies to 
changes in temperature rather than changes in pressure. A typical tempera- 
ture ratio will be obtained in the case of an aircraft traveling at high speed. 
The faster the aircraft flies at high Mach number, the hotter the outside 
surfaces become as a result of air friction and compression. The temperature 

ratio in this instance is between the temperature of the surrounding air or . 

the free air temperature, which is not affected by passage of the aircraft, and 
the temperature of the particles of air that are affected by contact with the 
aircraft. Within an engine, raising the pressure of the gas also raises the tem- 
perature of the gas. Similarly, expanding the gas lowers the temperature. 
These ratios are based upon absolute temperatures rather than upon Fahr- 
enheit or Celsius ( Centigrade) values. 

Total and Static Temperature and Pressure 

Imagine an aircraft in which very sensitive air pressure and temperature 
measuring devices are installed. If the aircraft is parked on the ground at 
an airport near the ocean and the instruments indicate an outside air tem- 
perature of 59°F and a barometric pressure of 29.92 in. Hg (standard sea 
level conditions), this would be the static ambient (or free air) temperature 
and pressure, respectively. 
When the aircraft takes off 
and flies over the ocean, 

: i—= 7 [ : — -— ----Temperature 59° 
Just above the water, the ere ria Altimeter Setting J 
Static temperature and a sa ag 

pressure have not changed 7 
and instruments which 
could sample these static 
values in flight would re- 
cord values identical to 
those read while on the 
ground. Let us assume, 
however, that the two sen- 
sitive instruments on the : onion eamen 7 
aircraft are designed to es ee ae 
measure the “total” pres- fle. 
Sure and temperature 
which the aircraft “feels,” 

as shown in Figure 3-4. Figure 3—4. Temperature and Pressure Rise 

Due to the Effect of Ram 
Because the aircraft is in 

flight, these instruments now indicate that both the temperature and pres- 
sure are higher than when the aircraft was Standing at the airport before 
flight. This increase over the static values of temperature and pressure is the 
ram effect. The amount of ram effect is a function of the airspeed. The two 
instruments are still feeling the static values which they indicated while the 
aircraft was not moving. Additionally, they now also feel the effect of the 
forward motion of the airplane. This is the ram temperature and ram pres- 



sure shown by the curves in the insert on the Flight Spectrum in Section I. 
The combination of static plus ram temperature or pressure is called total 
temperature or pressure. Although an aircraft in flight was used as an exam- 
ple, the same characteristics will be found wherever there is motion of a 
gas or liquid, whether it is against the outside of an object or inside of a 
machine or container, such as within the casing of a turbojet engine. 

In gas turbine engine analysis, references to the type of pressure or tempera- 
ture will be indicated by the subscript, “s” or “t.” Thus, T,, means the total 
temperature at Engine Station 7, and P,, means static pressure at Station 4. 
Both total and static values are used in various engine performance calcula- 
tions because one or the other will apply more specifically to particular char- 
acteristics, depending upon the circumstances. 

Delta and Theta 

When reciprocating engines are considered, performance corrections have 
to be made for changes in carburetor air temperature and pressure. These 
corrections are usually written out so that everyone may easily see what they 
are. However, in the case of gas turbine engines, which are much more sensi- 
tive than reciprocating engines to changes in engine inlet air temperature, the 
older correction method is too time-consuming. For gas turbines, the Greek 
letters, delta and theta, representing standard correction factors, are used to 
correct observed data to provide a standard for the comparison of engines 
operated under different conditions of temperature and pressure. These cor- 
rections apply throughout the entire engine, as well as at the engine air inlet. 
Rpm, fuel flow and exhaust gas temperature, for instance, are all corrected 
for a change in the state of the air entering the front of the engine. The 
values for delta and theta may be obtained from a table, or calculated: 

Observed absolute pressure 
29.92" Hg (or 2116 pst 

Observed Temperature Rankine 
= Theta = Observed Temperature (“Rankine 

Examples of the use of delta and theta for correcting engine performance 
will be shown later. 

& = Delta = 


In day-to-day performance corrections and analysis, there will be very little 
need for a knowledge of cycles, as long as the physical aspects of the proc- 
esses are known. A cycle is simply a process that begins with known condi- 
tions and ends at about the same conditions with which it began. For in- 
stance, a reciprocating engine cylinder on the intake stroke inhales a fuel/air 
mixture, the temperature and pressure of which have been raised by the 
engine supercharger somewhat over that of the surrounding air. Once the 
mixture gets into the cylinder, it is further compressed (cylinder volume 
decreasing), burned to raise its temperature and pressure (almost constant 





volume) and then expanded to produce work at the propeller shaft (volume 
increasing in the cylinder). The exhaust gases are then discharged through 
a port into the atmosphere from which the air originally came. The cylinder 
portion of this process is called a closed cycle because the cylinder receives 
a fresh charge of fuel/air mixture each time the gases are exhausted, and 
the process is begun anew. The supercharger, however, is working on an open, 
or continuous-flow, cycle identical to that of gas turbine engines. 

In the continuous-flow cycle of the gas turbine, the compressor raises the pres- 
Sure and temperature of the air taken in at the inlet. Fuel is introduced into 
the airstream and then burned in the combustion chambers to raise the tem- 
perature further. The entire mass of the gases is expanded through the tur- 
bines and the jet exhaust nozzle, all in one continuous process. Although 
the atmospheric conditions are slightly different at the entrance and the exit 
of turbine engines, the differences depending on airspeed, a cycle analysis 

(Reciprocating Engine Cylinder Only) 




Figure 3—5. Engine Operating Cycles 

can be made effectively. Graphic examples of these cycles are shown in Fig- 
ure 3-5. The important features of these cycles are that the several parts 
of the closed cycle (for reciprocating engines) take place entirely within 
one chamber and the cycle is constantly being repeated, while each part of 
the open cycle (for gas turbine engines) takes place in a different part of 
the engine and is a continuous process as long as the engine is running. 

Reynolds Number Effects 

When air in motion (or the movement of a body through air) is studied, 
there are three principal factors to consider: the velocity, the density and 


ee TT ee eee Se 

—_— —~ = 



the viscosity of the air. The effect of the changing viscosity does not become 
pronounced until the air density drops to a relatively low value. At the air- 
speeds normal for a turbojet aircraft, this low air density is usually encoun- 
tered at altitudes of approximately 35,000 feet, and above. At these high 
altitudes, the adhesive force between the individual air molecules (called 
viscosity) and the inertial (ram) force of a cubic inch of molecules in motion 
are both reduced below the normal sea level value for any given airspeed. 
However, the inertial force is reduced more, in proportion, than the viscosity. 
The result is that the boundary layer of air, which lies next to any body such 
as an aircraft wing or compressor blade in an airstream, begins to become 
appreciably thicker at and above, roughly, 35,000 feet. The molecules rush- 
ing by a wing or compressor blade surface are not able to tear as many other 
molecules off the surface as they could at sea level. Consequently, the drag 
and friction losses of the wing or blade begin to increase noticeably above an 
altitude of roughly 35,000 feet. This, in turn, reduces the operating effi- 
ciency of the aircraft or engine involved. The effect becomes greater as 
altitude is increased. This effect is ex- 
pressed by a factor called Reynolds 
Number Index, which is applied to 
the performance values obtained at or 
near sea level in order to correct the 
sea level performance to that which 
actually will be obtained at high alti- .ow 
tudes. Figure 3-6 illustrates the effect Low REYNOLDS NUMBER 
of Reynolds number on the efficiency Figure 3—6. Variation of Wing or 
of a wing or compressor blade. Compressor Blade Efficiency 
with Reynolds Number 




Mach Number 

There is a characteristic of gases which makes it very difficult for a certain 
velocity to be exceeded, either when the gas, itself, is in motion or when an 
object is moving through the gas. A few examples of this “restricting” effect 
would be the flow of a gas through a nozzle, an aircraft in flight, and the 
movement of a sound wave through air or some other medium. All of these 
reach a velocity above which it is difficult or impossible to accelerate further. 
In the case of sound, this gas characteristic actually determines the maximum 
speed at which a pressure wave or a warning signal can be transmitted up- 
stream, Because the speed of sound 
is the most common evidence of 
this characteristic, as well as the 
most easily measured, it is used as 
the basis for determining Mach 
number, which is the ratio of the 
speed of an object to the speed of 
sound in the same medium and at 




- 100 7 ‘ et eer Ge 

! a 
SS caben he Sina < 100 ae 14 the same temperature. The restrict- 

ing tendency is dependent only on 

Figure 3—7. Variation of Speed the temperature of the air and not 
of Sound with Temperature its density. The relation of Mach 




number to airspeed is shown on the Flight Spectrum. Mach number also 
varies with different kinds of gases, liquids and solids, as shown in F igure 
3-7. Mach number has been found to be a very convenient parameter in high- 
velocity flow work and as a measure of speed for high-performance aircraft. 

Basic Thermodynamics 

Thermodynamics is the study of heat flow and heat exchange. The subject 
involves the change from one heat level to another, either up or down the 
heat range, as well as the exchange of heat for mechanical work and mechan- 
ical work for heat. 

Letters and Symbols 

Certain principal quantities used in the Study or analysis of gas turbine en- 
gine thermodynamic processes are designated by letters and symbols, the 
most frequently used being as follows: 

A — Cross-sectional area SHP — Shaft horsepower 
c — Velocity of sound in air TSFC — Thrust specific fuel 
Cp, Cy — Specific heats at constant consumption 

pressure and volume 
C — Coefficient or constant 
ESHP — Equivalent shaft horse- 

T — Absolute temperature, °R, °K 
t — Gage temperature, a 

power (turboprop) V — Velocity 
ESFC — Equivalent specific fuel v — Volume 
consumption (turboprop) w— Weight 

F — Thrust 

g — Acceleration due to gravity, and 
mass conversion factor, 32.174 

w — Rate of flow 

§ — Delta—trelative absolute 

re ee haga A — Delta—difference or change 
M — Mach number—V/c » — Eta—efficiency 
N — Rotational speed, rpm 6 — Theta—relative absolute 
P — Absolute pressure temperature 
p — Gage pressure p — Rho—density 

Laws Governing Energy Exchanges 

Certain well-known laws have been found to govern energy exchanges. The 
basic ones are as follows: 

First Law of Thermodynamics — When mechanical work is transferred into 
heat or heat transferred into work, the amount of work is always equivalent 
to the quantity of heat. 

Second Law of Thermodynamics — It jis impossible by any continuous, 

self-sustaining process for heat to be transferred from a colder body to a 
hotter body. 


r PWA OPER. |NSTR. 200 

on Boyle’s Law — At constant temperature, the volume of a given weight of any 
gas varies inversely as the pressure to which the gas is subjected. 

Charles’ Law — At constant pressure, the volume of a given weight of any 
gas varies directly with the absolute temperature. 

Two other rules are used in gas turbine performance analysis, the first of 
which depends on the First Law of Thermodynamics. They are: 

f 1. The heat content of the gas at any point in the engine is equal to the 

| heat content of the entering air plus any additions and minus any ex- 

f tractions or losses ahead of the point being studied. This is sometimes 
called the Law of the Conservation of Energy. 

2. Everything that goes into the engine must come out again, as long as 
the engine is operating normally. This is sometimes called the Principle 

of Continuity. 

From these basic ideas comes a host of formulas, long and short, simple and 
complicated. It is not the function of this book to trace the basic ideas 
through this forest of formulas to the fairly simple working arithmetical 
calculations. The aim, here, is to show how day-to-day corrections and 
calculations can be made, in order to help those who must check an aircraft 
engine’s performance from time to time. 

How the Gas Turbine Works 


sin atee 

engine was given under How the Gas Turbine Operates. It was shown that 
the air enters the front of the engine, rises in temperature as the pressure in- 
creases within the compressor, and increases in temperature and heat content 
in the combustion chamber. The gases expand through the turbines to pro- 
duce the work needed to drive the compressors. The gases are then ducted 
rearward and out of the engine by the jet nozzle. 

j In Section I, a brief description of the thermodynamic actions within the 

=p || 

The work done on the air by the compressor is approximately equal to the 
mass airflow, times the increase in air temperature, times a constant. .The 
heat added in the combustion chamber can also be called work and equals 
the number of BTU added per unit of time divided by a constant. In the 
turbines, the work done by the hot gases is almost exactly the reverse of the 
work required by the compressor (plus the shaft horsepower supplied to the 
propeller in the case of a turboprop). The work in the turbine equals mass 
(weight) flow, times temperature drop, times a constant. In the jet nozzle, 
the work accomplished equals mass flow, times velocity squared, times a 
constant. During steady state operation, the work done by the turbines must 
be almost exactly the same as the work needed by the compressor (plus the 
‘4 propeller for a turboprop engine). 

—~»y = 

During engine acceleration, the turbines must provide excess driving power 
rTm~> to increase the rpm of the compressor and turbine rotor system. Likewise, 
during deceleration, the power output of the turbines must be less than that 





needed by the compressor (and propeller for a turboprop). Although the 
temperature drop available to obtain work from the very hot gases passing 
through the turbines is much greater than the actual temperature rise through 
the compressor, the mass (or weight) flow available to the turbines is exactly 
the same as the mass (or weight) handled by the compressor, assuming that 
the air bled from the engine is equal in weight to the fuel added. Therefore, 
a temperature drop takes place through the turbines that is roughly equivalent 
to the temperature rise which 
occurred in the compressor. 300 
However, Figure 3-8 shows 
that to obtain equal tempera- 
ture rises and drops beginning 
at 519°R (points a to b) in 
the compressor and at 2000°R 
in the turbine ( points c to d), 
the pressure change needed is 
much less at the higher tem- 
perature. This difference 
(points d to e) is why pressure 
is left to produce thrust in the 
turbojet. In the turboprop, 
more energy is taken out of 
the turbines to drive the propeller, with a small amount of pressure left over 
to induce jet thrust (points f to g). This characteristic, then, is the reason 
why turbojet engines are able to provide a forward thrust and why turbo- 
props and shaft turbines can produce large amounts of power at the propeller 
or drive shaft. The energy added in the form of fuel has been more than 
enough to drive the compressor. The energy remaining produces the thrust 
or power for useful work. 




400 800 1200 1600 2000 

Figure 3—8. Compression and Expansion 
Curve for Gas Turbine Engines 

Gas Turbine Performance 

One of the items of greatest interest to flignt crews, maintenance personnel 
and ground test stand operators is whether or not the engines are operating 
normally and producing rated thrust or power. Whenever circumstances 
warrant, data taken from the engine instruments must be corrected to stand- 
ard day conditions at sea level and compared with data known to be correct 
(or average data) for an engine of an identical type, series and model. While 
a certain amount of standard or normal data for comparison purposes will 
be supplied by the aircraft or engine manufacturer, each operator is en- 
couraged to correct the data observed during engine operation and to use 
the corrected data to establish normal average performance. After all of the 
data for enough normally functioning engines to furnish a reliable average 
have been corrected to standard conditions, deviation from the average by 




an individual engine will be readily apparent. In addition to furnishing a 
key to the location of a possible malfunction of an individual engine, these 
corrected data will also be useful (by making comparisons) in determining 
errors that may be present in the indicating equipment, where a gage OF other 
‘nstrumentation malfunction might otherwise lead the flight or maintenance 
crew to believe that something was wrong with the engine, itself. 

The following discussions and sample calculations are presented to illustrate 
the procedure used to correct observed gas turbine engine performance data 
to standard day conditions for use during both test stand and in-flight air- 
craft operation. The analysis for both the turbojet and turboprop engines 
includes actual numbers for hypothetical engines operating under represent- 
ative conditions. Three examples are shown for each engine: on the test 
stand, in the aircraft on the ground (static), and in the aircraft in flight. 
The test stand methods shown conform with Pratt & Whitney Aircraft pro- 
duction engine practice. In the examples shown for a dual axial compressor 
turbojet engine and for a single axial compressor turboprop engine, instru- 
ment readings peculiar to each engine type are noted. 

An aircraft gas turbine engine generally recognizes such variables as altitude 
and airspeed only by their effects on the temperatures and pressures occurring 
at the air inlet and gas outlet of the engine. For this reason, it is necessary 
to convert altitude, airspeed and similar data to temperatures and pressures 
before they can be used directly for engine performance calculations. 

The values for atmospheric pressure and temperature, aircraft speed and 
aircraft inlet duct pressure recovery in the following examples have been 
chosen as typical for a multiengine transport-type aircraft using gas turbine 
engines. Engine airbleed and accessory power extraction have been shown 
in these examples in order to make the normal correction procedure com- 
plete. Other values for other cases may be substituted, as desired. Following 
the numerical examples and the accompanying curves, certain internal engine 
characteristics are explained progressively through the engine. The general 
curves and data required to make the necessary corrections to observed en- 
gine performance data are also included. 

It should be noted that the engine data shown here cannot be used to analyze 
performance on specific engines. These examples are intended only to show 
methods used, since most of the data are entirely hypothetical. Pratt & 
Whitney Aircraft Specific Operating Instructions and Installation Handbooks, 
or the applicable aircraft Flight Handbook or military Technical Order, 
should be consulted for specific engine data pertaining to a particular engine 
series and model, whenever it is necessary to adapt the procedures illustrated 
here to actual practice in the field. The general tables and curves at the end 
of this section may be used for performance analysis, however. 






Basic Engine Analysis - Turbojet 

The following analysis illustrates the manner in which engine performance 
may be calculated from the data which are usually available from the air- 
craft instruments. Figure 3-9, below, reviews the engine station designations for 
a dual axial compressor tur- 
bojet engine. The curves and 
tables used in making many 
of these calculations are ref- 
erenced, as required, and may 
be found in this section. The 
engine curves presented are 

—_ @® 

3 25 

a > now 

53 fs 

@p | LowPressure | < 5 
cq Compressor | + Burner 
ams 2 


Pe = ust Duct 

Figure 3—9. Turbojet Station Designations 

for Analysis 


Operating Conditions 

. True barometer or alt pressure (Pam) 

in. Hg abs 

. Ambient air temp-static (Tam) °F 

. Ram pressure ratio (Pt2/Pam) 

. Ram temperature ratio (Tt2/Tam) 

. Engine inlet total press. (Pi2) in. Hg abs 
- Engine inlet total temp (Ttz) °R 
. Pressure correction factor (dt2) 

. Temperature correction factor (Ot2) 

. Square root of temp corr. factor (VO) 

Net thrust (F,) Ib 


Sea level 
ground test 
stand, bell- 
mouth inlet, 
standard jet 

(2070 psf) 

75 (535°R) 



535 (75°F) 






only, and are not to be used 
for actual calculations. 


Engine in- 
stalled in air- 
craft being 
checked stat- 
ically at sea 
level. Inlet 

duct recovery, 

95%. No ex- 
haust losses. 

(2150 psf) 

42 (502°R) 


502 (42°F) 




Aircraft in 
flight at 
30,000’, 500 
knots TAS. 

No inlet duct 

loss. No ex- 
haust duct 

(628 psf) 

—20 (440°R) 


500 (40°F) 



Ground ba- 
rometer or 
flight altim- 
eter. (Use 
alt tables, 
Fig. 3-24) 


For in-flight 
data, use 
Fig. 3-26 

For in-flight 
data, use 
Fig. 3-27 

® x® 
G) /29.92 
© °R/519 

Measured on 
test stand. 
See Note 5 



“=a — a ee ee ee ee 

— = cE 








ys Se 













Corrected net thrust (F:/dt2) lb 
Thrust correction for losses lb 
Fully corrected net thrust (Fi/dt2) Ib 
Fuel flow (wr) lb/hr 
Corrected fuel flow (ws/dea Vr) lb/hr 
Fuel flow corr. for losses lb/hr 

Fully corrected fuel flow (Wr/dts VO 12) 

Fully corrected thrust specific fuel con- 
sumption (TSFC/V0:) Ib/Ib/hr 

Exhaust gas temperature (Tt7) °R 
Corrected EGT (Ttz/®0t2) °R 
High pressure compressor speed (Nz) rpm 

Corrected high pressure compressor speed 
(N2/ V8 tz) rpm 

Turbine outlet total pressure (Pr7) 
in. Hg abs 

Engine pressure ratio (Pt:/Prz) 

Jet nozzle expansion ratio (Pt7/Pam) 

For Calculation of Engine Thrust in Aircraft 

Low pressure compressor speed (Ni) rpm 

Corrected low pressure compressor speed 
(N:/ Ve t2) rpm 

Ratio of specific heats (y) 

Gross thrust parameter (¥) 

Ratio of hot to cold nozzle areas 
(A hot) 
(A cold) 

Gross thrust coefficient (Cg) 

Nozzle area cold (A cold) sq ft 
Gross thrust (Fg) Ib 
Airflow parameter (Wa t2/dt2) Ib/sec 








































See Note 6 
& Item 




See Note 6 
& Item 

@ + © 


@ /® 


a) /® 


@ /® 
@ /@ 


@ /® 


See Note 5 

Use oy) on 

Fig. 3-12 





1. Use absolute temperatures for all temperature calculations. 
2. Altitude pressure at 30,000 feet is from altitude tables, Figure 3-24. 
3. Exhaust gas temperature (Item (9) is in degrees Rankine (°F + 460). 

4. N, is_noted only because it is used to determine the airflow parameter 
(Item 3.4)) on P&WA dual axial compressor turbojet engines. Other en- 
gines may use other airflow indicators. The specific engine Installation Hand- 
book or equivalent publication should be consulted for this information. 

5. Net thrust equals gross thrust at zero airspeed. For calculation of net 
thrust in flight, the following formulas are used: 

Ww, V Ot2 S12 Vv 
Net thrust (F_) = thrust (F_.) — § ———— x— x2 
(a) Net thrust ( 1) = gross thrust ( e) ( 5., Va. : ) 

Where: VO (Item G64) is read from curves on Figure 3-12 

V, = true airspeed in feet per second 
& = mass conversion factor = 32.174 

A hot 

(b) Gross thrust (F,) a ee C, X A cold xX ec 

-@ «Ox @x@x@ 

P,, Must be in pounds per square foot. All areas are in square feet. V,, is in 
feet per second. Psi ( ys) is gross thrust parameter (Item(29). 

6. Calculations of thrust and fuel flow losses resulting from airbleed, acces- 
sory power extraction and inlet and exhaust duct losses are as follows: 

(a) Thrust losses (AF,) — 

P,.-P P,.-P ae ae es Ww 
F =F tl ~ t2 + C ti * t2 4 Se +C bl ) 
sl. | P,, ‘ ( P. Ps i LPC 

Wi hp ext hp ext 
us cu ( W, ™ ¥ ox (32 ref Be * Cum (z ref ) onl 


F. = Corrected net thrust ( Item(1 }) 

y., = Total pressure at entrance of inlet duct 
Fi = Total pressure at compressor inlet 

F = Total pressure at turbine outlet 

Pe = Total pressure at exhaust nozzle inlet 

Wy = Mass of air bled from compressor, Ib/sec 
Ww, = Corrected engine airflow, Ib/sec 

hp ext = Horsepower extracted from the accessory drives 

*Ptz-Pis/Piz is to be used only when a nonstandard or extended nozzle is 



















hp ref = A reference horsepower, constant for a given engine. Value for 
the engine of this example is 1232. 

C,, Cy, C,x = Correction factors due to duct loss, airbleed and power ex- 
traction. These are shown on Figures 3-14 and 3-15, to be 
read at the proper value of corrected thrust. 

LPC = low pressure compressor 

HPC = high pressure compressor 

The fully corrected net thrust is the corrected net thrust (F,/8,.) plus the 


The assumed thrust losses connected with each case are as follows: 

Inlet duct loss 5% zero See Operating 
Conditions ,Case II 
Corrected true airspeed (TAS/\V@am) zero 543 TAS/ V @) /519 
Low compressor airbleed lb/sec 0.50 1.20 Measured or 
High compressor airbleed lb/sec NO 0.75 0.75 Measured or 
Low compressor power extraction hp LOSSES zero 25 Measured or 
High compressor power extraction hp ASSUMED 20 40 Measured or 
Duct loss correction factor (Ca) 0.58 — Use Fi/dam and 
with Fig. 3-15 
Airbleed corr. factor, LPC (C1) 1.20 1.92 Use F,/dam and 
with Fig. 3-14 
Airbleed corr. factor, HPC (Co:) 2.56 3.48 Use Fi/dam and 
Gé) with Fig. 3-14 
Power extraction factor, LPC (Cpx) 0.11 0.09 Use Fi/dam and 
with Fig. 3-14 
Power extraction factor, HPC (Cpx) 0.04 0.012 Use F:/dam and 
Gé) with Fig. 3-14 
Combined thrust loss ( AF») lb 754 220 Substitute the 

(b) Fuel flow losses (Aw,) — 

P..- : 
Aw, = W, | 7== +0; (737+) + ©, ( 
tl tl 

CG ( Wii ) =a G2 (2 <) er ee, ( 
w, JHPC hp ref JL PC 



hp ext 
hp ref 

above values in 
the previous 

) ave | 












Inlet duct loss (same as 63) 5% zero See Operating 
Corrected true airspeed (same as 69) zero 543 TAS/ VY (2) /519 
Duct loss correction factor (C,.’) NO zero zero Usually zero 
Airbleed correction factor, LPC (C,,’) LOSSES 0.40 0.38 Use Fi/dam and 
G8) with Fig. 3-14 
Airbleed correction factor, HPC (Cy:’) ASSUMED 0.91 0.78 Use F,/dam and 
with Fig. 3-14 
Power extraction factor, LPC (C,,’) 0.0385 0.025 Use F,/dam and 
with Fig. 3-14 
Power extraction factor, HPC (C,;’) —0.118 —0.097 Use F,/dam and J 
with Fig. 3-14 
Combined fuel flow loss (Aw;) lb/hr 290 20 Substitute the 

Ww, = corrected fuel flow (Item (©) 

C,', Cy’, C,.’ = correction factors, applying only to fuel flow and not to 
thrust, shown in Figure 3-14. 


All other quantities are identical to those used in thrust loss calculations. 
The fully corrected fuel flow is the corrected fuel flow plus the losses. 

The assumed losses connected with each case are as follows: 

above values and 
the losses from 
Note 6(b) in the 
above formula. 

7. The results of a test stand calibration are plotted on Figure 3-10 for a 
hypothetical turbojet engine. Since the quantities plotted are properly cor- 
rected (as shown), these curves can also be used to check the engine in the 
aircraft, as long as the altitude is below eines 35,000 feet. The only 
items affecting these curves are: 

(a) Ram pressure ratio. In flight, the net thrust obtained at a given engine 
pressure ratio decreases as the ram pressure ratio resulting from airspeed in- 
creases, as shown on Figure 3-10. The thrust, only, changes. A schedule of 
the thrust change can be shown for various values of ram pressure ratio (true 
airspeed), if desired. 

(b) Reynolds number effects decrease the efficiency of the engine above 
approximately 35,000 feet. These effects are not shown on this curve. 

(c) The curve shown is for the basic engine, without airplane installation 
effects. The curves can be redrawn to account for these effects, if desired. 

(d) A change in free air temperature merely changes the point on the 
curves at which the engine operates when at full throttle. If the air is colder 




than standard temperature (519°R), the EPR will be higher at any altitude. 
If the air is warmer, EPR will be lower. In either case, the engine will still 
be operating on these lines. | 

(e) Figure 3-10 is for one individual engine. Although data scatter will be 
obtained when a series of similar engines is plotted on the same curve, this 
will be an advantage because it will show the average value of each quantity 

obtained on a fleet of engines. 

(f) The full throttle (maximum thrust) point is not shown because this 
point will vary with engine trim. Actual values must be obtained from the 
Specific Operating Instructions or from other engine performance data which 
may be available. 

8. Figure 3-25 is a Temperature Conversion Table for converting Fahrenheit 
temperatures to Celsius (Centigrade) temperatures and vice versa. Such a 
table will prove quite useful when engine performance is calculated because it 
is necessary to be consistent in the temperature scale used throughout the 
calculations. Engine ground test stands are frequently instrumented for Fahr- 
enheit temperature, whereas use of the Celsius (Centigrade) scale is common 
practice for cockpit instrumentation in the aircraft. 



Figure 3—10. Corrected Turbojet 
Performance Curves for a 
Hypothetical Dual Axial 

Compressor Engine 



‘<< & 36-22 24 22 28.2) Se 








28 5000 6000 8000 

20 24 3.2 36 40 48 

Figure 3—11. Gross Thrust Coefficient | Figure 3—12. Engine Airflow Parameter 


SCALE "c" 
4 5 6 7 8 9 10 iN 12 
i 4. os ef lined er Se ae | 

SEBERY 499 4) 

fe \ | 13 
4.0 a 2 

. —— SCALE Mat 

> | eS 
a — 4 ai a { 
id ; ~~ 
a i — a oo 
j= ig / 
- 8. i as ——+ = 
2 1.3.4 > a : i 
a 1.30 ; : 
. O-yY 2 eee ees ee ee a ee = 
{ } 3 
e | a 
no i +—f——t-_} Qe 
> | Po ' 
x } a 
= w 2S 
ra tes ie A pe EE Ge BP 
= = 
0 Lan a 
7) ; 7) 
° ! ; oad 
: + + t (8 ww 
- : ° 
: « 

1.5 1.6 1.7 ——SCALE "a" 
SCALE"B"——= 2.0 2.5 3.0 3.5 4.0 4.5 


Figure 3—13. Gross Thrust Parameter 3 





r 4 






Figure 3— 

14. Estimated Engine Corrections for Airbleed and Power Extraction 









4000 8000 12000 16000 

Figure 3—15. Estimated Thrust Correction Factor Due to Duct Loss 

Basic Engine Analysis - Turboprop 

Internally, the turboprop differs from the turbojet chiefly in that it uses most 
of the energy of the hot gases to drive the turbine wheels which, in turn, drive 
the engine compressor and the aircraft propeller. This leaves only a sinall 
percentage (10-20 per cent) of the energy of the gases available for jet 
propulsion. A schematic diagram is shown below in Figure 3-16 to indicate 
station numbers. Following this are the sample calculations for a single axial 
compressor turboprop engine. 

a3 3 
a5 < 13 
5 3 6 E 
8 e Compressor Burner rtd a 
am | 2 3 4 5 7 

Figure 3—16. Turboprop Station Designations for Analysis 


. Engine inlet pressure (Pt2) 

. Cell static pressure (Pse) 


Operating Conditions 

_ True barometer pressure (Pam) in. Hg abs 

in. Hg abs 

in. Hg abs 

. Pressure correction factor (dt2) 

5. Pressure correction factor (5am) 


. Square root of temp corr. factor (V2) 
. Torquemeter pressure 
. Torquemeter constant 

. Ram pressure ratio (P:/Pam) 
. Ambient air temperature (Tam) °F 
. Engine inlet total temperature (Tz) °R 

. Temperature correction factor (42) 


. RPM observed 
. RPM corrected 

obs RPM 

Shaft horsepower observed (SHP) 


Sea level 
ground test 
standard inlet 
and exhaust 
systems. No 
power extrac- 
tion, duct 
loss, or air- 
bleed. Power 






71 (531°R) 

533: (75°F) 




Engine in- 
stalled in air- 
craft. Sea 
level static 
ground check. 
2% exhaust 
duct total 
pressure loss. 
4% inlet duct 
loss. Com- 
pressor air- 
bleed, 0.5 Ib/ 
sec. Power 
extraction, 40 
hp. Power 



85 (545°R) 

547 (87°F) 







check at 
25,000 ft of 
engine in air- 
craft. 2% ex- 
haust duct 
total pressure 
loss. 4% inlet 
duct loss. 
airbleed, 0.5 
lb/sec. Power 
extraction, 40 
hp. Power 
setting: Nor- 
mal Rated. 
True airspeed, 
400 knots. 






477 (17°F) 

a Be 




Use barom- 
eter reading 
for Cases I 
and II and 
Fig. 3-24 
for Case III. 
Use instru- 
mented Pts 
for Case I. 
Use Fig. 
3-24 for 
Case II. 
Use Fig. 
3-26 and(1) 
for Case III. 





For in-flight 

data, use Fig. 


reading (See 
Note 5) 





See Note 1 




2 & 



yx B 


ky & 




SHP corrected for Tam (SHP’) 

SHP corrected for Pam (test stand only) 
Engine mass airflow (w,) Ib/sec 
Airflow corrected for inlet duct loss 
SHP installation losses (ASHP) 

Jet thrust (F,) (test stand only) lb 
Jet thrust (F,) (in flight) Ib 

Engine pressure ratio (test stand only) 

SHP correction factor for EPR (test stand 

Fuel flow (observed) pph 

Fuel flow corrected for Tam (wr’) pph 

Fuel flow correction for airbleed (Aw:) 

Final corrected fuel flow (we) pph 

Final corrected SHP ( SHP”) 

Net jet thrust (F,”) Ib 

Jet horsepower 

Equivalent SHP 
SFC | 

The following is a summary of the corrections made above to account for 



losses and for nonstandard temperature and pressure. 

RPM correction 
SHP correction for Tam 
Airflow correction 

SHP installation losses (duct loss, air- 
bleed ) 

Horsepower extraction 

Fuel flow correction for Tee 

Fuel flow correction for installation losses 










See Note 3 
See Note 3 
See Note 6 
See Note 9 
See Note 9 
See Note 15 

See Notes 15 

and 16 
See Note 7 
See Note 11 

See Note 12 
See Note 13 

See Note 14 
See Note 17 
See Note 18 

Case I, i) 

y Ee 
Cases II and 

III, see Note 

@9) plus Gi) 
@8 / 
Q8) / 6D 


See Note 2 
See Note 9 
See Note 9 

Measured or 
from aircraft 

er’s data 

See Note 12 
See Note 13 

1. SHP = actual engine rpm x torquemeter pressure x K, where K is the 
torquemeter constant, which varies for each model of engine. The actual 
value of K depends upon the torquemeter calibration—psi or in. Hg. This 
formula is used in test stand calculations and for computing shp in a ground 
check of the engine in the aircraft. Calculation of corrected shp in flight re- 
quires the use of shp specification curves, appropriate power setting, and TAS. 







—— — — = TE 

— oD 


— on 




2. Corrections applied to calculations made from instrument data taken 
from the aircraft must be added to the calculation from observed data. In 
Cases II and III, Item (5) , observed shp delivered to the propeller shaft is 
a net figure. Losses due to power extraction and installation have already 
been extracted within the engine. It now becomes necessary to determine just 
what these were, in order to check engine output with specification curves 
that make no allowances for losses. 

3. In-flight values for shp (©) must be corrected for variation in T,,,, 
prior to any other corrections, by the use of the following equation: 

shp. A sh 
shp’ = shp 5) + | (=> @ std temp — ( 5 Ja Tom |B 

using A correction curves in Figure 3-17. 

4. Test stand correction of shp (5) is accomplished by use of the follow- 
ing equation: 

shp’ = shp (5) X —— 

where P,, is cell static pressure, in. Hg abs. 
5 In test stand calculations, no correction is made for T,,,. This correc- 

tion is incorporated in the basic turboprop engine curve, shp/8,. vs N/ V6 2s 
Figure 3-18. 
6. Engine airflow is determined from the turboprop specification curve 

(Figure 3-19) for appropriate conditions. If flight conditions prevail, use 
the following equations to determine T,, or TAS: 

2 ne 
T, ((R) = Tm (°R) + se. 640-4. -T.. 

aie Pig - Pre 
1. en = —— 

9, W,, = WwW, [: -*f | 

Using TAS, altitude, and shp’, enter the correction factor curve (Figure 
3-20) and determine C,” and C,,,”. These correction factors are then used in 
the following equation: 

A Fis A $ A Pz Ww 
= + C,” fb oe mw ( e1_ 
Ashp | P.. : ( P.. Pte ) 5 Cu ( Wa’ )| gd 
airbleed ( w,,) is given @ 0.5 Ib/sec 

inlet duct loss (AP,,/P,,.) 18 4% 
exhaust duct loss (AP,,/Pi,) is 2% 

w, is from step 


. JNSTR. 200 


a ie es 



shp’ is from step 

10. Jet thrust is corrected by use of the following equations: 

C, = 0.77 | 1 + 0.0524 “| 


C,, = 1.77 | 1 + 0.0524 “| 


A Pre A Pte A Pez | Wi 
AF, cate | P., + c,( P.- + Pre )+ Sas (= )| Ps @ Tete 

11. Applying EPR determined in Note 7 above, the shp correction factor, 

3-21. SHP’ x factor = shp”. 

12. Flight correction: 

Aw, = | ( — je i a (4")e . | “7 

Test stand correction: “, = 

13. Aw, = | (2) + C,,’ (=) | m © Ty 
tr a 

14. w.” = w, + Aw, 

, can be determined from the EPR correction factor curve in F igure 

15. Jet thrust and jet horsepower are read directly from the basic tur- 
boprop engine curve (Figure 3-18) as corrected values, the T,m Correction 
being incorporated in the curve. 

16. For flight calculations and ground check, read the turboprop spec- 
ification curve, Figure 3-19. 

17. shp” = shp’ + Ashp + hp ext 

is. FL” = F.’- AF. 

19. The values derived from the flight calculations may be further 
checked on the turboprop power setting curves (Figure 3-22 ). This curve 
is normally included in the Pratt & Whitney Aircraft Specific Operating In- 
structions for a specific engine model. A somewhat similar curve or table from 

which the same information may be obtained will also appear in the ap- 
plicable aircraft Flight Handbook. 

0. F, (ip) = <a 

This assumes a propeller efficiency of 80 per cent. 


X C 


sa HH FF as aa aH ua 





a 3000 
| 77) 
) O 
| * 
| aa 2000 
| = 
; ao 
co ey 1000 
35| 5 * 
f = alw 1000 
aes eee 
© goo w|é 
i a 
rs) 6) 

| en —1000 

| 3000 




LB / HR 


-80 -40 0 40 80 120 


Figure 3—17. Ambient Temperatufe Effect on Shaft Horsepower, 
Net Thrust and Fuel Flow 



(Lb/ Hr/SHP) 
(°F) tO000 
1500 J 
JET Fy 1100 
( Hp) 
10,000 200 400 600 800 11,000 200 400 600 
Figure 3—18. Basic Turboprop Engine Curves — 

aaa awe Ft Tweet agwa wat wai wea eS aw uaz 



(Lb) 500 

- 1.45 
2200 Fa = 
e . 
2000 ox 
(Lb/Hr) 1800 a 130 
1600 © 1.25 
40 # 1.20 
Wa 35 o 115 
(Lb/Sec ) a 
30 WwW 1.10 


4000 1000 2000 3000 4000 5000 6000 

Figure 3—21. Shaft Horsepower 
Correction for Engine 

% 100 200 300 400 500 600 Pressure Ratio 

Figure 3—19. Turboprop Specification Curve 


. 2000 4000 6000 1000 2000 3000 

Figure 3—20. Turboprop Shaft Horsepower Correction Factors 


) e ) 

do °%t1 — 3UNLVYSdWAL 

Ov + Oc+ 0 O02- Ov - 

43- 30NLILIW 3YNSS3Yd 
000‘0€ 000'02 000‘OI! TS 


31avMON1V XV 

Figure 3—22. Typical Turboprop Power Setting Curves 

Q3193dxX3 JTLLOYHL fav S¢ 



PWA OPER. | NSTR. 200 

Component Performance Details 

Some of the engine components require special consideration. The discussions 
which follow pertain to certain characteristics of the various sections of the 
engine, which must be dealt with separately when engine performance is 
calculated. Except for the discussions on the jet nozzle and the afterburner, 
these considerations are applicable to both turbojet and turboprop engines. 

Engine Air Inlet Duct 

Although the engine air inlet duct is not a “performing” part of the engine, 
it affects the engine thrust or power output by causing some inlet air pres- 
sure losses or distortions, which, in turn, affect the thrust of a turbojet 
engine or the shaft horsepower and jet thrust of a turboprop. The engine 
compressor inlet section is of no concern, since all calibration and testing is 
done with the inlet unit installed on the engine. However, the effects of the 
duct or fairing, supplied by the aircraft manufacturer to conduct the outside 
air into the engine, must be accounted for when engine performance is calcu- 
lated. While pressure losses vary for different conditions in the inlet duct, 
such as those caused by the flight attitude of the aircraft, the full ram tem- 
perature rise is always realized. 

Because the air passing through a bellmouth inlet on a turbojet engine has 
very little pressure loss and because the pressure rise (or loss) is low when 
the engine is on the test stand, the values used in Case I of the turbojet anal- 
ysis are reasonably correct. However, when either a turbojet or turboprop en- 
gine is operated on a test stand, an inlet pressure survey should first be made 
to determine the corrections, if any, which should be taken into consideration 
for the test stand equipment. The barometer reading must be corrected to 
account for any losses or gains at the inlet. 


Although the type of turbojet engine used for the sample calculations is a 
dual axial compressor model, the methods shown are applicable to either 
dual or single axial or centrifugal compressors. Likewise, although the turbo- 
prop analysis is for a single axial compressor engine, the same system of cal- 
culations can be used for dual compressor turboprops, or other types, by 
making slight changes. The difference between turbojet and turboprop engine 
calculations lies chiefly in the temperature and pressure differentials within 
each type of engine and in the means used to check over-all engine perform- 
ance. Thrust is used for the turbojet, and torque (or horsepower) is used for 
the turboprop. 

Because mismatching of the two units of a dual axial compressor engine is 
rare in a properly designed gas turbine engine, the two compressor units 
are treated as a single compressor. Should any mismatching occur, it will 
be indicated by a significant variation in such quantities as the ratio of N, 
to N, and will require a more detailed performance analysis than is normally 
used. To analyze the performance of mismatched compressors properly 
would require more information than can be given in this book. 





The function of a compressor is to increase the density of air going to the 
burners. Some losses are incurred during compression, principally because 
a compressor cannot operate at 100 per cent efficiency. Consequently, the 
normal rise in compressor air temperature is accompanied by an additional 
temperature rise, due to the inefficiency of the compressor. The additional 
rise in compressor temperature also results in a pressure rise lower than would 
be obtained in a perfect compressor. 

Burners and Turbines 

Although the engine instrumentation on most gas turbine-powered aircraft 
is adequate for proper engine control, the normal instrumentation found on 
the average cockpit instrument panel is not sufficiently complete to enable 
an individual analysis to be made of either burner or turbine performance. 
Whenever necessary, the foregoing engine performance calculations may be 
supplemented to assist in pinpointing a difficulty within the engine. Burner 
pressure can be obtained by instrumenting the burner pressure sensing line 
to the fuel control and can be used as a ratio (P»/Pt2 and P,/P:z) to deter- 
mine where an engine difficulty is located, when necessary. It should be noted 
that the burner pressure used to schedule the fuel control on most Pratt & 
Whitney Aircraft engines is a static and not a total pressure. Both the com- 
pressor inlet and turbine outlet pressures normally are obtained as total pres- 
sures, so that the ratios found between these and burner pressure provide only 
working values and not absolute values of pressure ratio. 

Jet Nozzle 

Normally available instrumentation provides all necessary data for calcu- 
lating and analyzing the performance of the jet nozzle. The nozzle area 
can be measured, the turbine outlet temperature and pressure are known, the 
atmospheric pressure is easily obtained and curves showing nozzle coefficients 
are available. Since most Pratt & Whitney Aircraft turbine engines have fixed 
nozzle areas, including afterburning engines with two-position nozzles, the 
problem of determining jet nozzle performance is simplified. Not enough is 
presently known regarding the effects on nozzle areas and coefficients of ex- 
haust silencers and thrust reversers for their inclusion in these calculations. 
Jet thrust performance of turboprop engines is not normally checked during 
routine testing. If shaft horsepower and fuel consumption are satisfactory, 
the jet thrust is also assumed to be satisfactory. It is further assumed that 
jet thrust is directly related to the shaft horsepower being developed. 


The afterburner on Pratt & Whitney Aircraft turbojet engines is so designed 
that the basic engine thrust will be approximately the same whether the 
afterburner is off or on, when two-position (fixed) nozzles are used. There 
are occasional variations in engine pressure ratio or turbine discharge pres- 
sure, in exhaust gas temperature and in compressor rpm. These are small, 
however, and will not be dealt with here. The over-all effect of the after- 
burner is to increase the temperature and, therefore, the velocity of the gases 



on Si A A lB pa 8 


Se = 

i a ag sll ian aad lea aaa - 

—- — Sinai 

ars a 


ee — = 

ee eee 

_ wm mT eT aw aa == @&e 






discharged through the nozzle, 
thereby increasing the thrust. 
An increase in thrust of 50 per 
13,000 cent, or more, is available on 
LB a sea level test stand. Several 
p00 times this amount is available 
at high flight speeds with a 
two-position (fixed or nonad- 
justable area) nozzle. A typical 
fixed-nozzle (two-position) 
afterburner check curve, Fig- 
ure 3-23, is included for use in 
conjunction with plotting tur- 
bojet performance, Figure 
3-10. To use this curve, the 
test stand must be capable of 
1.80 measuring thrust and total 
(basic engine plus afterburner ) 
fuel flow. When the operation 
of the basic engine is satisfac- 
tory, the engine performance, 
with afterburning, should be 
approximately as shown in Fig- 
ure 3-23. No correction of aft- 
Figure 3—23. Afterburner Check Curve erburner readings is necessary. 











-20 fe) 20 40 60 80 100 120 


The thrust and power losses from driving accessories are relatively small 
because the ratio of the power required to drive the compressor to the power 
extracted for driving the accessories is usually low. However, these losses 
are present and must be accounted for when engines are checked. In some 
cases, such as with the high pressure compressor-turbine unit of a dual axial 
compressor turbojet engine, for instance, extracting power to drive accesso- 
ries actually increases the thrust output of the engine. This also raises the 
turbine inlet temperature. The normal procedure to control inlet temperature 
properly is to reduce the thrust or shaft horsepower by an amount com- 
mensurate with the power being extracted. 


When an accessory is driven by compressor airbleed, the loss of thrust (or 
shaft horsepower in a turboprop) is much greater than when the same acces- 
sory is driven mechanically. Also, airbleed used for cockpit and cabin 
pressurizing or air conditioning is expensive in terms of both thrust (or 
shaft horsepower) and fuel consumption. In deciding the manner in which 
accessories will be driven, the aircraft manufacturer must weigh the cost 
in operational losses against the additional weight which would be incurred 
if auxiliary power units were used to drive the accessories and air condi- 



PWA OPER. |[NSTR. 200 

tioning unit. Some losses also result when airbleed is used to provide anti- 
icing heat for either the engine or the aircraft. All of the losses resulting 
from the use of airbleed as a source of power or heat must be taken into 
consideration when engine performance is calculated. It also must be re- 
membered that the use of airbleed, especially during full-throttle engine 
operation, will raise the turbine inlet temperature to some extent, which will 
probably be evidenced by a rise in exhaust gas temperature. 

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Figure 3—24. Standard Altitude Tables 



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INSTR. 200 



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Figure 3—27. Compressor Inlet Temperature vs True Airspeed 

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Section IV _ 4}——i 

Gas Turbine Engine Operation — 

The engine operating information in this section is necessarily both general 
and abbreviated. Pratt & Whitney Aircraft’s General Operating Instructions 
are available for those who must pursue the subject in more detail. The Gen- 
eral Operating Instructions provide a comprehensive background and expla- 
nation of the engine operating characteristics with which both maintenance 
and flight personnel should be familiar to operate an aircraft powered by a 
particular engine type. General Operating Instructions also contain engine 
operational procedures which will serve as a guide to those who use the in- 
structions for training. P&WA General Op- 
erating Instructions No. 190 are applicable 
to dual axial compressor commercial and 
military nonafterburning turbojet engines 
of the JT3 (J57), JT4 (J75) and JT8 
(J52) series. P&WA General Operating 
Instructions No. 191 are applicable to dual 
axial compressor military afterburning tur- 
bojet engines of the J52, J57 and J75 
series. Before attempting to operate any 
engine, the appropriate P&WA Specific 
Operating Instructions, aircraft manufac- 
turer’s Flight Handbook or military Tech- 
nical Order should be consulted for spe- 
cific operating data and procedures. 

The most outstanding feature of a gas turbine engine is its simplicity. The 
total lack of intricate valve and timing mechanisms, complicated reciprocat- 
ing piston arrangements or literally thousands of moving parts leads to re- 
markable reliability and ease of operation. The pilot monitors comparatively 
few gages, signals the amount of thrust or power he desires to the fuel control 
(which does most of his computing for him) and that is all that there is to it. 


. JNSTR. 200 




Once the desired engine thrust or power condition is set up, the fuel control 
will maintain an approximately constant per cent of thrust or power output 
with a fixed throttle position, compensating for changing conditions at the 
compressor air inlet. Even if a malfunction should develop, the engine will 
continue to operate and produce propulsive force under most conditions. 

Engine Ratings 

Centrifugal and single axial compressor engines 
usually develop their full rated thrust (or horse- 
power, in the case of a turboprop) at 100 per 
cent rpm. This is sometimes referred to as “speci- 
fication thrust.” Because the thrust or power of 
engines of these types is approximately propor- 


Gy WwW NY 

tional to rpm, lower thrust or power ratings such INDICATES 

as “Normal Rated” are developed at specific rpm F, (thrust) 
values below 100 per cent. These lower ratings ON CENTRIFUGAL 
are obtained in flight or on the ground by adjust- AND SINGLE AXIAL 
ing the throttle to a setting at which the rpm indi- COMPRESSOR 
cated by the aircraft tachometer will be the same ENGINES 

as that specified for a given engine rating by the 
engine manufacturer. The rpm values for each en- 
gine rating are stipulated in the appropriate En- 
gine Operating Instructions, Flight Handbook or military Technical Order. 

The pilot obtains the various engine ratings in a different manner on dual 
axial compressor engines. Although the discussion which follows pertains 
specifically to engine thrust ratings for dual axial compressor engines, the 
definitions given for these ratings also apply generally to ratings for gas tur- 
bine engines of other types. Rpm cannot be used directly to set or check en- 
gine thrust on dual axial compressor turbojet engines because full rated thrust 
is usually obtained at some rpm value below 100 per cent on the tachometer. 
Furthermore, as will be explained in detail later, full rated thrust is not ob- 
tained at the same rpm on individual engines of the same model and series. 

P&WA dual axial compressor turbojet engines are rated according to their 
thrust. The full thrust ratings are normally based upon the thrust obtained 
with the throttle in the full forward position under sea level static standard day 
conditions. On some engines, espe- 
cially commercial models whose rat- 
ings may be more conservative, full 
rated thrust may be obtained at throt- 
tle positions below the full forward 
position. The ratings must be corrected 
by the use of charts or curves if they 
are to apply to other conditions of 
ambient temperature or barometric 
pressure. The ratings below full thrust, 
COMPRESSOR such as Maximum Continuous or Nor- 
ENGINES mal Rated, are based upon allowable, 
but not instrumented, turbine inlet 

Fr (thrust) 


a TT eT Gi 




temperature for either a specific period of time or continuous operation, as 
the case may be. 

All of the engine ratings except those of Maximum, Take-off (wet or dry ) 
and Military are obtained by adjusting the throttle to the position at which 
the fuel control will maintain the allowable turbine inlet temperature in ac- 
cordance with its predetermined schedule. Although the fuel control will 
maintain an approximate percentage of Take-off (dry), Military or Normal 
Rated thrust for moderate variations in conditions at the compressor inlet, 
this does not mean that turbine discharge pressure or engine pressure ratio 
will also remain constant. It is necessary, therefore, to refer to curves or charts 
furnished by the aircraft manufacturer or published in the aircraft Flight 
Handbook in order to determine the turbine discharge pressure or engine 
pressure ratio which will represent the thrust condition or rating desired by 
the pilot. These curves or charts will be adaptations, corrected for a partic- 
ular aircraft installation, of the curves discussed later. Once the value for 
turbine discharge pressure or engine pressure ratio is determined for the 
prevailing compressor inlet temperature and altitude, the throttle is adjusted 
to obtain the computed value on the turbine discharge pressure or engine 
pressure ratio gage. 

The commercial and military ratings of the engines are not the same and are 
described separately. When applying these ratings, it is necessary to use the 
correct group of ratings. The two methods of rating the engines are necessary 
because of the difference in commercial and military engine requirements. In 
military aircraft, the urgency of the mission frequently determines the manner 
in which the engine is to be operated. When the engine is installed in com- 
mercial passenger or cargo aircraft, the time between engine overhauls and 
maximum reliability are of primary concern, and more conservative engine oper- 
ation becomes the rule. Consequently, military versions of the engines might 
be described as “full thrust” engines, whereas commercial versions, in many 
cases, are “part thrust” engines. That is, rated thrust is obtained at less than 
full throttle position. The commercial “Take-off (wet)” and military “Take- 
off” ratings are comparable. The same applies to commercial “Take-off (dry)” 
and military “Military.” “Maximum Continuous” for commercial engines 1s 
equivalent to “Normal Rated” for military engines. There is no military coun- 
terpart for the commercial “Normal Rated” and “Maximum Cruise” engine 
ratings, nor is there a commercial counterpart of the military “Maximum.” 


Take-off (wet) — This is the maximum allowable thrust for take-off. The 
rating is obtained by actuating the water injection system and setting the com- 
puted “wet” thrust with the throttle. The rating is restricted to take-off, may 
have an altitude limitation, and is time-limited. Engines without water in- 
jection do not have this rating. 

Take-off (dry) — This is the maximum allowable thrust without the use of 
water injection. The rating is obtained either by placing the throttle in the 
full forward position or by adjusting the throttle to the computed Take-off 
(dry) thrust for the existing ambient conditions, depending upon the manner 




in which the engine is rated. The rating is time-limited and is to be used for 
take-off only. 

Maximum Continuous — This rating is the maximum thrust which may be 
used continuously, and is intended only for emergency use at the discretion 
of the pilot. The rating is obtained by adjusting the throttle to a predetermined 
turbine discharge pressure or engine pressure ratio. 

Normal Rated — Normal Rated thrust is the maximum thrust approved for 
normal climb. The rating is obtained in the same manner as Maximum Con- 
tinuous. On some engines, Maximum Continuous and Normal Rated thrust 
are the same. 

Maximum Cruise — This is the maximum thrust approved for cruising. It 
is obtained in the same manner as Maximum Continuous. 

Idle — This is not an engine rating, but, rather, a throttle position suitable 
for minimum thrust operation on the ground or in flight. It is obtained by 
placing the throttle in the Idle detent on the throttle quadrant. 


Maximum (afterburning engines only) — This rating is obtained by placing 
the throttle in the full forward position and actuating the afterburner. The 
rating is time-limited. 

Take-off (nonafterburning engines only) — This rating is obtained by actu- 
ating the water injection system and placing the throttle in the full forward 
position. The rating is restricted to take-off, may have an altitude limitation, 
and is time-limited. Engines without water injection do not have this rating. 

Military — This rating is obtained by placing the throttle in the full forward 
position without the use of water injection. The rating is time-limited to 30 
minutes. This rating serves for take-off in aircraft powered with engines that 
do not have water injection. 

Normal Rated — This rating is obtained by placing the throttle in a position 
selected by reference to turbine discharge pressure or engine pressure ratio 
and associated curves. Normal Rated thrust and all lower flight thrust selec- 
tions may be used continuously. 

Idle — Idle thrust is obtained by placing the throttle in the Idle detent on 
the throttle quadrant. 

Engine Operating Variable (Dual Axial 
Compressor Engines) 

Thrust, as such, cannot be measured directly on an installed turbojet engine. 
Therefore, some other engine variable such as compressor rpm, turbine dis- 
charge pressure, or engine pressure ratio, all of which vary with thrust, must 
be employed as an indication of the propulsive force which the engine is 
developing. Turbine discharge pressure or engine pressure ratio, rather than 
compressor rpm, is normally used as the operating variable for measuring 



me. | eee | — ay —ay —y ~~4 —~y 


—_ 2 = Te 


PWA OPER. | NSTR. 200 

the thrust output of dual axial com- 
pressor turbojet engines. Engine pres- 
sure ratio gages, which are sometimes 
known as “ratiometers,” measure the 
air pressure as it enters the engine and 
the gas pressure at the turbine dis- 
charge. The ratio between the two 


(P,,/P,.) is read directly on the instru- NS” 
ment dial. This is the engine pres- 
sure ratio. : 

Single compressor turbojet engines commonly employ compressor rpm as an 
indication of engine thrust. However, many complications arise when rpm is 
utilized as the controlling variable for dual axial compressor engines. Some 
of the disadvantages are listed below: 

1. Since only the high pressure compressor rpm (N.) is governed by the 
fuel control, rpm does not provide an accurate means of determining 
whether or not the whole engine is functioning properly. 

2. Rpm for either or both rotors for any given thrust condition will vary 
slightly among individual engines, depending upon the engine trim speed. 
This variation in rpm must be taken into consideration whenever rpm is 
used to measure the thrust being developed by the engine. 

3. One per cent variation in rpm results in approximately four per cent 
variation in thrust at the higher thrust settings for the low pressure com- 
pressor rotor (N,) and five per cent variation for the high pressure com- 
pressor rotor (N,), whereas one per cent variation in turbine discharge 
pressure or engine pressure ratio results in only one and one half per cent 
variation in thrust. 

4. Rpm for either or both rotors does not vary in direct proportion to the 
thrust being produced by the engine over the entire thrust range. 

For these reasons, Pratt & Whitney Aircraft recommends that turbine dis- 
charge pressure or engine pressure ratio be used as the engine variable for 
indicating thrust on dual axial compressor engines. The use of either of these 
is not only much simpler than the use of rpm for dual axial compressor en- 
gines but is considerably more accurate as well. 

Because the thrust developed by the engine is indicated by the difference in 
pressure or the pressure ratio (which is also proportional to engine thrust) 
between the engine air inlet and the discharge pressure at the jet nozzle, 
turbine discharge pressure, by itself, should not be used directly as an accurate 
indication of engine output. Compressor inlet pressure (P,,) must also be 
taken into consideration on the curves or in the charts used whenever turbine 
discharge pressure, alone, is instrumented on the aircraft. P,, will be the true 
barometric pressure (corrected for duct loss) for a static engine or a com- 
puted value based on airspeed and pressure altitude for an aircraft in flight. 
Engine pressure ratio gages, on the other hand, consider compressor inlet 
pressure automatically for all conditions. 




Engine Instrumentation 

Although engine installations may differ, depending upon the type of both 
the aircraft and the engine, gas turbine engine control will usually be obtained 
by the use of the following instrumentation. 

One of the following instruments, depending upon the installation, may be 
provided to indicate the thrust which the engine is developing. Reference to 
both types of pressure gage is made in the text of this book because, at the 
present writing, either gage may be found in use. Of the two, the turbine dis- 
charge pressure gage is usually the most accurate gage, due, primarily, to its 
simplicity. The gage may be installed either permanently on the aircraft or, in 
some instances, temporarily, such as during an engine trim. An engine pres- 
sure ratio gage, on the other hand, is simpler for the pilot to use because it 
automatically compensates for airspeed and altitude by taking compressor 
inlet pressure into consideration. As the accuracy of the pressure ratio gage 
continues to improve with further development, it is anticipated that the in- 
strument may become standard for most dual axial compressor turbojet en- 
gine installations. 

Thrust — Turbine Discharge Pressure Gage — This gage indicates the total 
engine internal pressure immediately aft of the third stage turbine (P,,), and 
serves as an indication of the pressure available to generate thrust, when used 
in conjunction with compressor inlet pressure ( gy 2 

Thrust — Engine Pressure Ratio Gage — This gage indicates the engine 
pressure ratio as a measure of the thrust being developed by the engine. Be- 
cause the compressor inlet may not be instrumented directly for P,., the P,, 
sense for the pressure ratio gage must be placed at some other location on 
the aircraft, preferably as near the engine air inlet as possible. The gage, 
therefore, will not exactly sense the true compressor inlet pressure. For accu- 
rate thrust indications, the difference between the instrumented pressure and 
the true value for P,, must be taken into consideration, either by an applied 
correction factor or by a built-in correction within the gage itself. 

Torquemeter (Turboprop Engines) — Because only a small 
part of the propulsive force is derived from the jet thrust, 
neither turbine discharge pressure nor engine pressure ratio 
is used as an indication of the power produced by a turbo- 
prop engine. Turboprops are usually fitted with a torquemeter, 
operated by a torquemeter ring gear in the engine nose section, similar to the 

torquemeter provided on large reciprocating engines. The torque being de- | 

veloped by the engine is proportional to the horsepower. Torquemeter oil 
pressure is used to indicate shaft horsepower (SHP). 

Engine Speed — Gas turbine engine speed is measured by 
the compressor rpm which, of course, will also be the turbine 
rpm. Tachometers are usually calibrated in per cent rpm so 
that various types of engines may be operated on the same 
basis of comparison. On centrifugal and single axial compressor 




turbojet engines, compressor rpm serves directly as an indication of the en- 
gine thrust being produced. It has been stated that this is not a recommended 
procedure for dual axial compressor engines, the principal purpose of the 
tachometer being for use during engine starting and to indicate an overspeed 
condition, should one occur. On dual axial compressor engines, the high 
pressure compressor is the one most frequently instrumented, although, par- 
ticularly on test installations, the low pressure compressor may also be fitted 
with a tachometer. The value of 100 per cent rpm does not necessarily in- 
dicate full engine thrust for engines of this type. Due to manufacturing toler- 
ances, it is impossible to divide the compressor work precisely between the 
two compressors. Therefore, the high pressure compressor speed is adjusted 
by means of trimming the engine, as will be explained later, until the engine 
produces its full rated thrust. The rpm which results is the engine trim 
speed. This speed is stamped on the engine data plate. The actual rpm at 
which the tachometer will read 100 per cent depends upon the tachometer 
gear ratio of a given engine model and may be obtained from the P&WA 
Specific Operating Instructions or the aircraft Flight Handbook. 

Exhaust Gas Temperature— Exhaust gas temperature (EGT), 
tailpipe temperature or, more specifically, turbine discharge 
temperature are one and the same thing. This temperature 
is an engine operating limit and is used to monitor the 
mechanical integrity of the turbines as well as to check engine 
operating conditions. Actually, the temperature at the turbine inlet is the 
important consideration, this being the most critical of all of the engine 
variables. However, as has been pointed out, it is impractical to measure 
turbine inlet temperature in most engines, large models especially. Con- 
sequently, temperature thermocouples are inserted at the turbine discharge 
instead, this temperature providing a relative indication of that at the inlet. 
Although the temperature at this point is much lower than that at the inlet, 
it enables the pilot to maintain surveillance over engine internal operating 
conditions. Several thermocouples are usually used, spaced at intervals around 
the perimeter of the engine exhaust duct near the turbine exit. The exhaust 
gas temperature gage in the cockpit indicates the average of the temperatures 
measured by the individual thermocouples. Because the importance of exhaust 
gas temperature cannot be overemphasized, the subject is discussed in more 
detail later. 

Fuel Flow — Fuel flow gages indicate the fuel flow in pounds 
per hour from the engine fuel control. Fuel flow is of funda- 
mental interest to flight crews in monitoring flight fuel con- 
sumption, checking engine performance, and in setting up 
flight cruise control. 


10 e 
~— FUEL FLOW wy 

Engine Oil Pressure — To guard against engine failures re- 
sulting from inadequate lubrication and cooling of the various 
engine parts, the oil supply to critical areas must be moni- 
tored. The oil pressure gage usually indicates the engine oil 
pump discharge pressure. 





Engine Oil Inlet Temperature — The ability of the engine oil 
to perform its job of lubricating and cooling is a function of 
the temperature of the oil as well as the amount of oil sup- 
plied to the critical areas. An oil inlet temperature gage is 
frequently provided to indicate the temperature of the oil as 
it enters the oil pressure pump. Oil inlet temperature also serves as an indi- 
cation of proper operation of the engine oil cooler. 

Fuel Pump Inlet Pressure — Fuel system characteristics fre- 
quently make it advisable to monitor the fuel pump inlet 
pressure. In case of fuel flow stoppage in flight, it is desir- 
able to locate the source of the difficulty quickly, in order to 
determine whether it has occurred in the engine or in the 
aircraft fuel system so that prompt corrective action may be taken. In addi- 
tion, the fuel pump inlet pressure will indicate possible cavitation at the 
fuel pump inlet in flight and proper fuel system operation during engine 
ground checks. 

Engine Inlet Air Temperature — Thrust or power varies with 
the temperature of the air entering the engine at the com- 

decreases. The thrust which the engine will develop for any 
given ambient condition must, therefore, be computed for 
each inlet temperature. On most engines, particularly large ones, it is not 
considered advisable to instrument the compressor air inlet directly because 
of the possibility of a temperature probe failure with resultant damage to 
the engine’s internal parts. It is customary to measure the outside air tem- 
perature instead, correcting as may be necessary to obtain the true total 
temperature at the engine air inlet. Free (ambient) air temperature gages 
should indicate the total temperature, which includes the ram temperature 
rise, at some location as near to the air inlet duct as possible. 

The above represents the minimum instrumentation considered adequate for 
control of the engine. Some installations may have additional gages, numerous 
warning lights and the like to assist the pilot in operating his engine. Gen- 
erally speaking, present practice dictates restricting the number of instruments 
to only those that are really necessary. This enables the pilot and his crew to 
concentrate on operational features of the aircraft. 

Preflight and Starting 

During preflight inspection, particular attention should be paid to the engine 
air inlet, the visual condition and free movement of the compressor and tur- 
bine assembly, and to the parking ramp area fore and aft of the aircraft. The 
engine is started by means of an external power source or self-contained com- 
bustion starter unit. Starter types and the engine starting cycle have been 
discussed previously. On multiengine aircraft, one engine is usually started 
by use of a ground cart which supplies the air pressure for a pneumatic starter 
on the engine or fuel and air for a combustion starter when these are not 
carried aboard the aircraft. Airbleed from the first engine started is then used 
as a source of power for starting the other engines. 


pressor air inlet, the thrust increasing as the temperature 


During the start, it is necessary to monitor the tachometer, the oil pres- 
sure gage and the exhaust gas temperature gage. The normal starting se- 
quence is (1) to rotate the compressor with the starter, (2) to turn the igni- 
tion on, and (3) to open the engine fuel valve, either by moving the throttle 
to Idle or by turning a fuel shutoff lever or switch. Adherence to the procedure 
prescribed for a particular engine is necessary in the interest of safety and in 
order to avoid a hot or “hung” start. A successful start will first be noted by 
a rise in exhaust gas temperature. In the event that the engine does not 
light up within a prescribed period of time or if the exhaust gas starting 
temperature limit is exceeded, the entire starting sequence should be repeated 
again from the beginning. When necessary, the engine is cleared of trapped 
fuel or gases by continuing to rotate the compressor with the starter but with 
the ignition and fuel turned off. 


Most gas turbine engines do not require a prolonged warm-up, even in cold 
weather. The engine instruments and the various aircraft systems may be 
checked after the engines have been started and while the aircraft is taxiing 
to the runway take-off position, thus saving much valuable time. There is no 
need for the customary reciprocating-engine, part or full throttle engine run- 
up before take-off. Taxiing and engine operation at Idle should be held to a 
minimum, since turbine engine fuel consumption is relatively high during all 
ground operation. Taxiing too close to or directly behind other aircraft should 
be avoided. Their wake will invariably contain small stones and other trivia 
injurious to a hungry gas turbine engine. Once the aircraft is in position 
and ready for take-off, the brake is locked and the engine is run up at full 
throttle for a final instrument and thrust or power check, after which the 
brake is released and the take-off is made without further ado, using water 
injection or the afterburner, if desired and available. Take-off thrust or 
power on centrifugal or single axial compressor turbojet engines is checked 






by means of rpm and, in the case of a turboprop, torquemeter pressure. The 
procedure for checking a dual axial compressor engine for full (take-off) 
thrust is discussed later. 


Climb and Cruise 

Pratt & Whitney Aircraft Specific Op- 
erating Instructions contain altitude 
thrust setting curves similar to Figure 
4-1 for computing the throttle setting 
for varying conditions of altitude, com- 
pressor inlet total temperature and the 
required amount of engine thrust, 
which will depend upon the aircraft 
gross weight. These curves may appear 
in the Flight Handbook for a _partic- 
ular aircraft in a different form than 
that illustrated or the data may possi- 
bly be presented in a chart or table. 



Charts or curves furnished by the air- 
craft manufacturer which appear in 
the Flight Handbook are consulted, 
first, to determine the amount of thrust - 40 0 iat 
which will be required for the desired een eee ere ae 
airspeed and the known gross weight f ; 

of the aircraft. For climb this might pire f-1. Type Wrest Saien 

be, for instance, Normal Rated thrust Ceres Per Sains Peeeirs Tate 

or, possibly, 90% Normal Rated. The curves in the Specific Operating In- 
structions or those in the Flight Handbook are then used to determine the 
value for engine pressure ratio which will represent this amount of thrust 

for the existing ambient (outside) air temperature and altitude. The throttle 

is then adjusted to obtain the desired reading on the engine pressure ratio 
gage. Similar curves for setting power on a turboprop engine will be used 

for determining torquemeter pressure instead of engine pressure ratio. 

For military aircraft, either Normal 
Rated or higher thrust (or power) may 
be used for climb, depending upon the 
conditions encountered and the results 
desired. Referring to Figure 4-2, if the 
take-off is made at Point A, Point B 
will be reached at approximately the 
same time, whether Military or Normal 
Figure 4—2. Aircraft Climb Curve Rated is used for climb. Slightly less 
(Military Engine Ratings) fuel may be consumed when the climb 
is accomplished at the higher thrust 
setting. However, higher engine speeds and temperatures will also occur, re- 
sulting in slightly decreased engine life. Hence, it is recommended that Nor- 
mal Rated thrust be employed in military aircraft for normal climb. Y 




| 5 



= xD 




For commercial aircraft, Pratt & Whitney Aircraft recommends that either 
Normal Rated thrust or, preferably, a little less than Normal Rated thrust be 
used for climb. In either case, cruising altitude will be reached at approx- 
imately the same time. The amount of thrust used for climb is relatively un- 
important, provided it is in the vicinity of Normal Rated. If the climb is too 
slow, however, some loss of time will be experienced, even though no loss of 
range may result. If the aircraft is to cruise close to the approximate altitude 
for maximum range and cruising thrust, then, as a rule of thumb, approx- 
imately 10 per cent more thrust should be used for the climb than is intended 
to be used for cruising. 

In fighter-type aircraft, when a climb 
is made in afterburning, the desired 
altitude will be reached with the same 
(or possibly a little less) amount of 
fuel being used than if the climb were 
made without the use of the after- 
burner. This is due to the greatly re- 
duced time required. However, when 
climbing to cruising altitude on a cross- 
country flight, it is best that the after- 
burner not be used. Although cruising 
altitude would be reached with roughly 
the same amount of fuel by either 
method, the horizontal range or dis- 
tance obtained during the climb will 
be a great deal more when the after- 
burner is not used. This, in turn, 
will increase the maximum range of 
the aircraft. 

For cruising, Pratt & Whitney Aircraft 

recommends that thrust be established 

by setting the throttle to obtain a desired turbine discharge pressure or engine 
pressure ratio rather than setting the throttle to obtain the desirable combina- 
ion of altitude and Mach number. 

Once the selected, initial cruising altitude is reached, maximum range will 
be obtained by keeping the established Mach number constant, and allowing 
altitude to increase gradually as the aircraft gross weight decreases due to 
fuel consumed. The rate of climb will be very small; therefore, rather than 
endeavoring to climb at some specified rate, the altitude of the aircraft should 
be checked at frequent intervals to assure that the proper “climb flight path” 
is being maintained. The mission profile will vary according to changes in the 
outside air temperature. When an aircraft is flown on civil airways, the neces- 

sity to maintain a constant altitude flight plan may require that the altitude ~ 

be increased in steps at predetermined points along the route rather than at a 
gradual, constant rate of climb. There is no objection to this method of cruise 
control except that some loss of range will result. 

At a fixed throttle position, once the desired thrust condition has been set up 





for climb or cruise, the fuel control will maintain an approximately constant 
percentage of Normal Rated thrust (military) with varying compressor inlet 
conditions even though turbine discharge pressure will decrease as altitude is 
gained. Engine pressure ratio, on the other hand, which varies with compres- 
sor inlet temperature, will increase as the ambient temperature becomes lower 
at the higher altitudes. 

The use of fuel flow as a thrust-setting parameter will not be discussed, since 
curves for this purpose must be furnished by the aircraft manufacturer for a 
given aircraft and engine installation. It is important for the pilot to have and 
use curves or tables for computing fuel flow as, at the very least, the fuel flow 
gage may be employed to provide an “in the ball park” check on proper en- 
gine operation. Erratic or abnormal fuel flow will be the first indication of a 
malfunctioning engine or fuel control. 

In turbojet and turboprop aircraft, aircraft performance is closely allied to 
the skill, experience and degree of attention to detail of the pilot and his 
flight crew. The high cruising speed shortens the time available to diagnose 
an error and effect corrective action. Conversely, the dividends for operating 
at optimum performance are great. Rapidly changing gross loads, as a result 
of high rates of fuel flow, and the effects of varying altitude, airspeed and, 
especially, outside air temperature, combined with the economic premium to 
be obtained from precise operating technique, tend to concentrate the efforts 
of the turbojet or turboprop flight crew on striving for best performance 
rather than the mere handling of the aircraft and engine controls. As range 
requirements continue to increase for both military and commercial aircraft, 
greater emphasis must be placed on maintaining a high level of operating 
efficiency. Even the method of presenting flight data to the pilot, as it re- 
lates to the ease with which he will be able to interpret essential cruise 
control information, becomes a matter of great importance in gas turbine- 
powered aircraft. 

Letdown, Approach and Landing 

Gas turbine-powered aircraft 
are capable of very rapid rates 
of descent. Therefore, pene- 
tration and approach patterns 
are predicated upon this char- 
acteristic. Under instrument 
flight conditions, holding alti- 
tudes will be high. Once land- 
ing clearance has been ob- 
tained, letdown to the ap- 
proach traffic pattern will be 
accomplished as quickly as 
possible. Because turbine en- 
gines accelerate more slowly 
than reciprocating engines, it 



_ Re eae mR DP aD ae wa BE BaP ee Be a Be SS & 

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| 5 





is advisable to maintain as high an rpm as possible during the final approach 
in order to minimize the time required to attain full thrust should a go- 
around become necessary. Reverse thrust, when available, should be applied 
soon after landing while at high ground speed. With the turbojet, this is nec- 
essary to prevent recirculation of hot exhaust gases into the engine inlet. In 
the case of the turboprop, it is necessary because the higher the airspeed, the 
more effective is the reverse thrust. 

Engine Shutdown 

Engine shutdown on the ground is accomplished by cutting off the fuel to 
the engine and allowing the compressor to decelerate. Before the fuel is shut 
off, a brief cooling period at Idle, following high rpm engine operation, may 
or may not be required, depending upon the type of engine. After the fuel 
has been shut off, all engine valves and switches should be turned off, follow- 
ing the sequence prescribed in the applicable operating instructions. It is ad- 
visable to observe the engine closely during compressor deceleration, noting 
whether the compressor decelerates freely and making sure that there is no 
evidence of an internal engine fire. 

Engine Inlet Anti-icing 

In general, centrifugal compressor types of gas turbine engines are not as sus- 
ceptible as axial compressor engines to the formation of ice at the compressor 
inlet, although air inlet ducts may ice up regardless of engine type. Usually, 
unless the design of the duct is particularly prone to icing, the formation of 
wing ice on aircraft powered with centrifugal compressor engines will be a 
much more serious problem 
than engine icing and will be 
the limiting factor governing 
flight in icing conditions. Ax- 
ial compressor turbojet en- 
gines, on the other hand, are 
seriously affected by the same 
general conditions which 
cause wing ice to form. Ice 
accumulates on the compres- 
sor inlet struts and guide 
| : vanes, restricting the flow of 
= air into the engine. Fre- 

quently, ice in the air inlet 

duct will be as troublesome 
as ice on the struts and guide vanes. Unfortunately, the inlet duct (which is 
part of the aircraft and not the engine) is not as easily anti-iced as the engine 
compressor inlet. Under severe icing conditions, the inlet duct and the com- 
pressor inlet struts and guide vanes can ice up with startling rapidity. For a 
given icing condition, the compressor inlet of smaller engines with their 
closer inlet strut and guide vane spacing will be more susceptible to ice than 
larger engines. 





The rate of inlet icing in a given icing condition is relatively constant up to a 
moderate airspeed in the vicinity of 250 knots true airspeed. Above this speed, 
the rate of icing increases rapidly with increased airspeed. Because it affects 
the rate of air into the engine, the thrust or power setting also has a bearing 
on the amount of ice formation. This thrust or power effect, however, is 
offset by the fact that, even though the amount of icing may be greater at the 
higher throttle settings, more thrust or power from the engine will also supply 

more and hotter anti-icing air to eliminate the icing. The effect of airspeed is 

explained by the fact that at the lower airspeeds, air is sucked into the inlet 
duct, whereas at higher airspeeds it is rammed into the duct. During the suc- 
tion process, there is little, if any, change in the concentration of liquid water 
in the air entering the inlet and that of the outside air. When the air is rammed 
into the duct, most of the water droplets suspended in the air in the projected 
area ahead of the inlet duct tend to pass through the duct. Some of the air in 
this same projected area, however, is deflected around the duct. The inertia of 
the relatively heavy droplets causes them to leave the air in which they were 
suspended and they enter the duct instead of being included in the flow of 
deflected air. This has the effect of increasing the concentration of water 
entering the engine at high airspeeds. It also means that the rate of inlet icing 
will increase. Conversely, decreasing the airspeed below approximately 250 
knots true airspeed during icing conditions will reduce the rate of engine air 
inlet icing. 

The suction of air into the inlet duct increases with decreasing airspeed, be- 
coming greatest at zero airspeed or static conditions, when the rpm is high. 
The suction results in a drop in the static pressure of the air passing through 
the compressor inlet, especially in the vicinity of the inlet guide vanes. The de- 
creasing pressure at the inlet results in a drop in the temperature of the air, 
which may be as high as 10°F (approximately 6°C). This means that if 
visible moisture is present when an aircraft is operating on the ground or at 
low airspeeds and high rpm, such as during take-off, ice may form in the 
compressor inlet even though the ambient temperature is several degrees 
above freezing. 

Most axial turbojet and turboprop engines are provided with a means to tap 
hot, high pressure airbleed from the compressor and duct it through hollow 
inlet struts and compressor inlet guide vanes. Such a system warms the parts 
that are subject to icing, preventing the formation of ice. The system is anti- 
icing and not deicing. Therefore, icing conditions 
should be anticipated and, whenever possible, the 
anti-icing system should be turned on in advance of 

flying into icing conditions in order to warm up the Vee! o || WHEN IN 
engine inlet before ice actually commences to form. i (re) = a 
Usually, an anti-icing regulator is provided to regu- {y, pa 
late, automatically, the flow of anti-icing air with HEAT 

changing inlet temperature. Ice formation will be 
most severe when an aircraft is operating at low 
rpm and high airspeed, such as during a rapid 
descent. However, normally only intermittent use of high rpm during descent 

under severe icing conditions is necessary with Pratt & Whitney Aircraft dual” 


X C 



axial compressor engines. At the relatively high thrust settings used during 
climb or cruise, the anti-icing system will supply excess heat for protection 
against the accumulation of ice in the inlet section of the engine. To provide 
a desirable added margin of safety during operation in the lower thrust range 
for prolonged periods under severe icing conditions, the engine should oc- 
casionally be accelerated to higher thrust settings to provide excess heat for 
short periods of time. Ice formation at the engine inlet will most likely be 
noted by a rise in exhaust gas temperature, usually accompanied by a 
loss in thrust ‘or power, and possibly by compressor stall. Inlet ice may 
form whenever icing conditions are encountered, either in flight or during 
ground operation. 

Trimming Dual Axial Compressor Engines 

The practice of rating Pratt & Whitney Aircraft dual axial compressor turbo- 
jet engines according to their thrust results in an engine characteristic not 
common to most other turbojet engines. Due to manufacturing tolerances, 
each individual engine produces its rated or “specification” thrust at a slightly 
different compressor rpm. It has been pointed out that high pressure com- 
pressor (N:) rpm is governed by the fuel control on engines of this type. An 
adjustment on the fuel control enables this governed speed to be varied 
within specified limits. The adjusted high pressure compressor rpm, Cor- 
rected to standard day static conditions at sea level, is known as the “engine 
trim speed.” Engines are adjusted to produce their exact rated thrust on a 
calibrated engine test stand at the time of manufacture. The engine trim 
speed, or data plate speed as it is sometimes called, is then stamped on the 
engine data plate in terms of both actual rpm and per cent rpm. The fuel 
control adjustment is called “engine trimming.” Engines equipped for water 
injection will have two “trim speeds,” one for “dry” operation and one for 
“wet” operation when water injection is used. Pratt & Whitney Aircraft dual 
axial compressor engines may be subsequently retrimmed in the field from 
time to time when necessary, to compensate, within limits, for thrust de- 
terioration caused by foreign deposits on the compressor blades as the engine 
accumulates operating time. The fact that individual engines produce rated 
thrust at slightly different rpm is one of the reasons why rpm, as a definite 
numerical value, loses much of its significance as an engine operating variable. 

Military engines are adjusted or trimmed to produce Military or Take-off 
thrust at full throttle. Commercial engines, because they are, in effect, “part 
throttle” engines, are trimmed to produce their “dry” or “wet” rated thrust 
at some throttle position below the full forward position. Trimming is ac- 
complished by inserting a stop in the engine fuel control which will restrict 
throttle travel during the trimming process. Although commercial engine 
trim speed will therefore be lower than the high pressure compressor (Nz) 
rpm at which rated thrust will be developed by the engine, the trim speed 
will, nevertheless, be directly proportional to the rpm obtained at rated 
thrust. Data plates will be specific in indicating either the actual trim speed 
of an individual engine or the speed at which full rated thrust will be devel- 
oped. On commercial engines, neither the engine trim speed nor the speed 





obtained when the engine is developing full rated thrust will have any con- 
nection with the rpm which might be obtained were the engine to be op- 
erated at full throttle. Although military and commercial engines differ in 
this respect, the following discussion of trimming and speed bias applies 
to both. 

Proper engine trim is a prerequisite to proper engine operation. A thorough 
understanding of engine trim and trimming procedure is essential when dual 
axial compressor engines of this type are operated, since the use of turbine 
discharge pressure or engine pressure ratio as the primary measure of engine 
thrust and the concept of “speed bias” represent a departure from methods for 
trimming and operating other types of turbojet engines. To understand these 
engine operating characteristics better, it would be well to explain the basic 
philosophy adopted by Pratt & Whitney Aircraft in the design of the first JT3 
(J57) engine. This same philosophy has been perpetuated in the development 
of all subsequent Pratt & Whitney Aircraft dual axial compressor turbojet en- 
gines, such as the JT4 (J75) and the JT8 (J52). 

Three decisions were made during the initial development of the JT3, which 
resulted in the engine operating differently from previous turbojet engines of 
the centrifugal or single axial compressor type. 

1. Each engine was to produce rated or specification thrust exactly, at stan- 
dard day, sea level static conditions. This accounts for the variation in trim 
(data plate) speed among individual engines. 

2. Each engine was to operate at a constant turbine inlet temperature for a 
fixed throttle position. This is the reason for speed bias, as will be explained 
later. ° 

3. Each engine was to be built with a fixed exhaust nozzle area: This becomes 
important when turbine discharge pressure or engine pressure ratio is used 
as a measure of the engine thrust. | 

To determine the effect of these basic decisions, each will be discussed in detail. 

In order for each individual engine to develop its exact rated thrust at full 
throttle or at a given position of the throttle, it is necessary to compensate 
for the inherent variables among engines due to manufacturing tolerances 
on parts, varying friction loads, slight differences in compressor efficiency, 
and the like. One method for accomplishing this is to increase or decrease 
engine speed, as required, until the desired thrust is produced. This is the 
method used on Pratt & Whitney Aircraft dual axial compressor engines. 

Each engine is given a calibration run, at the time of manufacture, in a 
ground test stand which measures thrust directly. The engine is operated at 
some speed just above the rpm at which the intercompressor overboard air- 
bleeds close. After approximately ten minutes of operation to ensure that the 
engine has.completely stabilized, the thrust and the high pressure compressor 
(N,) rpm are recorded. The throttle is then advanced to some new position, 
and a second set of readings is taken. This is repeated until sufficient positions 
have been selected at approximately equal increments of thrust to cover the 
normal operating range of the engine. Each set of readings is then corrected 










to sea level, standard day conditions, as previously explained. A curve is con- 
structed by plotting corrected N., against corrected thrust. This curve repre- 
sents the performance of the individual engine in terms of rpm and thrust (in 
pounds). In the case of military engines, the curve is entered at the static 
thrust rating of the engine, corrected to standard day conditions at sea level, 
to determine the corrected rpm at which that particular engine will develop 
its rated thrust. In the case of commercial engines, the curve is entered at 
the thrust which should be developed by the engine when operating at the 
throttle setting determined by the engine trim stop inserted in the fuel con- 
trol. In either event, the rpm determined from the curve is the speed which 
is stamped on the engine data plate. Obviously, each engine will have its 
own calibration curve and its own data plate or trim speed. Only by rare 
coincidence will any two engines have the same speed on their data plates. 

As has been pointed out, corrected rpm and thrust values are used primarily 
for presenting data in terms of sea level, standard day conditions so that one 
engine may be compared with another or checked from day to day under 
nonstandard day conditions. These corrected values have no physical sig- 
nificance in themselves. Outside of determining what rpm should be stamped 
on the data plate, they have nothing to do with engine trimming procedures, 
subsequent engine thrust checks or take-off thrust settings. 

The decision to use turbine inlet temperature as the basic limit for engine 
operation is entirely logical, since this temperature (which represents the tem- 
perature to which the first-stage turbine blades are exposed) is probably the 
most important temperature within the engine. By adopting this philosophy, 
the engine is permitted to develop the maximum available thrust for any 
ambient (outside) air condition while, at the same time, operating tempera- 
tures within the engine are held to safe limits. 

On a hot day, the high pressure compressor (N.,) rpm for any given thrust 
condition will be higher than that on a standard day. On a cold day, the 
rpm will be lower than that on a standard day, 
for the same condition. This change in speed with 
varying compressor inlet temperature which, in 
turn, varies with the ambient air temperature, is 
known as “speed bias.” It is built into the fuel con- 
trol. The speed bias for any given engine model or 
series is represented by a Temperature-Rpm Curve, 
similar to Figure 4-3, which is published in the 
Pratt & Whitney Aircraft Specific Operating Instruc- 
tions and, normally, in the aircraft Flight Hand- 
book. The slope of the temperature-rpm (or speed bias) lines on the curve 
is determined solely by the N, change required to maintain a constant 
turbine inlet temperature. 

The use of speed bias (constant turbine inlet temperature) as a design pa- 
rameter results in an engine that develops more than rated thrust on a cold 
day and less than rated thrust on a hot day. However, the amount of varia- 
tion is not as great as that for a constant-speed engine. This change in 
thrust for a constant-speed engine may vary from 130 per cent of rated 













thrust at O°F to as low as 80 per cent 
at 100°F. On the other hand, the 
thrust of a dual axial compressor en- 
gine will vary only from approximately 
125 per cent of standard day thrust 
at O°F to around 85 per cent on a 
100°F day. This means that there will 
be more available thrust when it is 
sorely needed on a hot day, yet plenty 

-50 , *° of thrust will be retained on cold days. 

Figure 4—3. Typical Temperature- From this discussion, it should be obvi- 

RPM Curve ous that Pratt & Whitney Aircraft dual 
axial compressor engines are not constant rpm engines; neither are they con- 
stant thrust engines; nor do they operate to constant corrected rpm or thrust. 
They are essentially engines which operate at a constant turbine inlet tem- 
perature. The only time that military models of this type of engine will de- 
velop full rated thrust exactly at full throttle while operating at data plate 
or trim speed is on a standard day at sea level. Commercial models will 
develop a given fixed percentage of rated thrust only while operating at trim 
speed on a standard day at sea level. This explains the statement, previously 
made, that a fixed throttle position will maintain approximately a fixed per- 
centage of available thrust. Available thrust is defined as the thrust that a 
normally functioning engine will develop at the prevailing ambient conditions 
of temperature and barometric pressure. 

Since all engines of the same series and model are built with the same fixed 
exhaust nozzle diameter and area, the turbine discharge pressure (P,,) and 
engine pressure ratio (EPR) vary proportionally with the thrust produced on 
all Pratt & Whitney Aircraft dual axial compressor engines. For any given 
ambient air condition, every engine will produce the same P,. or EPR for 
the same thrust. Since the relationship between P,. or EPR and thrust is 
known, it is possible to construct a curve (called the Engine Trim Curve) for 
each model, which indicates P,. or EPR variation with compressor inlet tem- 
perature (T,.). When P,. is used as the engine operating variable, barometric 
pressure must also be taken into consideration. This is not necessary when 
EPR is used. The Engine Trim Curve appears in Pratt & Whitney Aircraft 
Maintenance Handbooks and applicable military Maintenance or Overhaul 
Technical Orders (Figure 4-4). A somewhat similar curve, except for the 
tolerances applied, also appears in the Pratt & Whitney Aircraft Specific 
Engine Operating Instructions or the aircraft Flight Handbook as the Take- 
off Thrust Check (or Setting) Curve. The Engine Trim Curve is used for 
engine trimming. Trim curves will also be provided for so-called “wet trim- 
ming” when water injection is used. The Take-off Thrust Check Curve or 
its equivalent in the aircraft Flight Handbook (which may be in tabular 
form) is used to check military engines prior to take-off to ensure that the 
engine is operating properly and is developing full Military thrust for the 
existing ambient conditions. Commercial engines will always be able to attain 
Take-off (dry or wet) thrust because these ratings will be obtained at less 




heh mH He HHP Hm HH Hw WH Fe BP aw aE ERB a es am ua 


than full throttle. Flight Hand- 
books for commercial aircraft 
will therefore contain Take-off 
Thrust Setting Curves because 
a full thrust check is neither pos- 
sible nor necessary. 




Both the Engine Trim and 
Take-off Thrust Check (or Set- 
ting) Curves must be corrected 
and redrawn by the aircraft 
manufacturer to reflect aircraft 
air inlet duct loss (sometimes 
called “ram recovery’), acces- 
sory drive loss and airbleed loss 
peculiar to any particular air- 
craft and engine installation. Curves published in aircraft Flight Handbooks 
or “dash one” series military Technical Orders can normally be expected to 
have these corrections incorporated. The basic Engine Trim Curve and the 
basic Take-off Thrust Check (or Setting) Curve furnished by Pratt & Whit- 
ney Aircraft or published in military Maintenance or Overhaul Technical 
Orders are applicable only to engines equipped with a bellmouth inlet and 
installed in a ground test stand. 


90 50 -50 0 +40 

Figure 4—4. Typical Take-off 
Thrust Check Curve 

When the Engine Trim Curve is used for retrimming, an engine can be 
trimmed in the field to compensate for any thrust deterioration which might 
result from accumulated deposits of dirt and/or scale on the compressor 
blades. Periodic retrimming in the field, as engine operating time builds up, 
can be accomplished only within specified rpm limits. If the full rated thrust of 
an engine cannot be restored without exceeding the allowable trimming lim- 
its stipulated in the applicable Pratt & Whitney Aircraft Maintenance Manual 
or military Technical Order, the engine should be “field-cleaned,” as explained 
later. It is never allowable to “up-trim” an engine to compensate for inlet duct 
loss, airbleed loss or accessory drive loss. To do so would result in an over- 
trimmed engine, operating considerably above the prescribed limits of tem- 
perature, rpm and thrust. 





It is essential to know how to compute and correct for speed bias when trim- 
ming an engine. It will be noted from Figure 4-3 that the vertical scale to the 
left of the Temperature-Rpm Curve is labeled “Per Cent Engine Trim or Data 
Plate Speed.” This is sometimes referred to as the “speed bias factor.” On some 
Temperature-Rpm Curves, this scale may be labeled, “Per Cent Standard Day 
Sea Level Static Military (or Take-off) Thrust,” which is just another way of 
expressing the same thing. Using the line which represents Military Rated 
thrust, for military engines, or Take-off (dry), for commercial engines, it is 
possible to enter the curve with the prevailing ambient temperature to de- 
termine the per cent of trim or data plate speed at which the engine will 
operate at the existing temperature. The actual tachometer reading (in per 
cent) at full throttle which should be observed if the engine is functioning 

a —— oe —s 

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properly is determined by multiplying the per cent obtained by the use of 
the curve by the actual trim or data plate speed (in per cent). By reversing 
the process and commencing with the observed tachometer (N,) reading 
and dividing by the per cent of trim or data plate speed (speed bias factor) 
obtained from the curve for the existing ambient temperature, engine trim 
speed (in per cent N,) may be determined. This will probably correspond 
exactly to the trim speed stamped on the engine data plate only if the en- 
gine has never been trimmed subsequent to the original calibration at the 
factory. In no case should the variance between the computed trim speed 
and the data plate speed exceed the maximum rpm adjustment allowed for 
retrimming. Should this be the case, it indicates that the engine has been 
overtrimmed, above permissible limits. It should be noted that this method 
of computing speed bias merely determines the equivalent or corresponding 
speed of the engine for a standard day and has nothing to do with the correc- 
tions to sea level, static standard day conditions, previously mentioned. The 
use of speed bias and the Temperature-Rpm Curve is only for the purpose 
of engine trim adjustment or when rpm is employed to check. engine per- 
formance and/or full thrust (military engines). 

The actual trimming of an engine, both with and 
without water injection, should be made in accord- 
ance with the procedures outlined in the applicable 
Pratt & Whitney Aircraft Maintenance or Overhaul 
Manual or the appropriate military Technical Order. 
In general, this procedure will consist of obtaining 
the ambient (outside) air temperature and the field USE CORRECT TEMPERATURE 
or true (not sea level) barometric pressure immedi- Sen ae ene 
ately preceding the trimming of the engine. If the 
engine is mounted in a test stand and fitted with 
a bellmouth inlet, compressor inlet temverature ( T,.) is usually instru- 
mented on the bellmouth screen and may be read directly. On an engine in- 
stalled in an aircraft, however, T,, is not instrumented. It is therefore neces- 
sary to use an ordinary thermometer to obtain the outside air temperature 
which, under static conditions, will be approximately the same as that at the 
compressor inlet. Care must be exercised to obtain a true temperature com- 
parable to that of the air which will enter the engine. This, for instance, 
would not be the temperature observed by the airport control tower and re- 
ported by radio, nor would it be a temperature measured in the glare of the 
hot sun. Neither should the control tower “altimeter setting” or the meteoro- 
logical station barometer reading be used to obtain the barometric pressure, 
since both of these will invariably be corrected to sea level. 

On the basis of this information, the desired turbine discharge pressure (P,.) 
or engine pressure ratio (EPR) is computed from the published Trim 
Curve, as called for by the applicable instructions. Even though an aircraft 
may be instrumented for engine pressure ratio (EPR), it is customary at the 
present writing to install a calibrated P,, gage temporarily whenever an engine 
trim run is to be made. (This is because the accuracy of present-day P,. gages 
is greater than the accuracy of EPR gages. As EPR gages continue to improve, 


u & 






it is anticipated that this type of gage may also be employed in the future for 
engine trimming. ) 

The engine is operated at full throttle (or at the fuel control trim stop) for 
a sufficient period of time to ensure that it has completely stabilized and 
that there is no thrust overshoot. Five minutes is the recommended sta- 
bilization period. A check should be made to ensure that the overboard air- 
bleed valves have fully closed and that all accessory drive airbleed for which 
the trim curve has not been corrected (such as, for instance, a cabin air- 
conditioning unit) has been turned off. 

When the engine has stabilized, a comparison is made of the observed and 
the computed turbine discharge pressure (or EPR) to determine the approx- 
imate amount of trimming required or, possibly, to ascertain whether the en- 
gine requires trimming. If a trim is necessary, the engine fuel control is then 
adjusted to obtain the target P,, or EPR on the gage. It should be noted that 
the engine is trimmed to P,, or EPR and not to rpm. Immediately following 
the fuel control adjustment, the N, tachometer reading is observed and re- 
corded. Fuel flow and exhaust gas temperature readings should also be taken. 

The observed N., tachometer reading is next corrected for speed bias by means 
of the Temperature-Rpm Curve. As previously pointed out, the observed 
tachometer reading is divided by the per cent trim speed obtained from the 
curve. The result is the new engine trim speed in per cent, corrected to stand- 
ard day (59°F or 15°C) temperature. The new trim speed in rpm may be 
calculated when the rpm at which the tachometer reads 100 per cent is known. 
This value may be obtained from the appropriate engine manual or Specific 
Operating Instructions. If all of these procedures have been performed satis- 
factorily, the engine has been properly trimmed. 

An alternate trimming procedure may be used on some military engines to 
avoid prolonged full throttle operation in the interest of reducing engine noise 
levels and high fuel consumption. The alternate method is a “part throttle” 
trim, usually at the 81.5° throttle travel position. It is particularly adaptable 
to nonafterburning engines and is similar to the fuel control trim stop method 
for commercial engines. Whenever this trimming method is used on an 
engine, an 81.5° throttle position curve on the Engine Trim Curve will be 
found in the appropriate Pratt & Whitney Aircraft Maintenance or Overhaul 
Manual or military Technical Order. 

Engine trimming should always be carried out under precisely controlled 
conditions with the aircraft headed into the wind. Precise control is neces- 
sary to ensure maintenance of a minimum thrust level upon which aircraft 
performance is based, to avoid needless engine problems in flight which might 
result in excessive exhaust gas temperatures, and to avoid possible engine stall 
during critical aircraft maneuvers. In addition, precise control of engine trim- 
ming contributes to better engine life in terms of both maximum time between 
overhaul and minimum out-of-commission time due to engine maintenance 
requirements. Engines should never be trimmed under icing conditions. 





Cleaning Engine Air Passages 

Field-cleaning of the air passages in the compressor of an axial compressor 
turbojet or turboprop engine may be accomplished by introducing a clean- 
ing agent at the engine air inlet while the engine is operating. The process 
is intended for use on engines displaying definite evidence of performance 
deterioration due to the accumulation of deposits of foreign material on the 
compressor vanes and blades. On dual axial compressor engines, perform- 
ance deterioration may be detected by the repeated necessity to increase the 
engine trim speed to maintain the full rated thrust of the engine. An engine 
should be field-cleaned whenever full thrust cannot be restored by field trim- 
ming without exceeding the allowable “up-trim” rpm limits. If the full rated 

engine thrust is not recovered after field-cleaning, the engine must be re- 

moved from the aircraft and sent to overhaul. 

{4 —\ 

The only material approved for field-cleaning Pratt & Whitney Aircraft 
gas turbine engines is a cleaning agent known as “Carboblast—Jet Engine 
Type” (Military Specification MIL-B-5634, Blasting Grit, Soft). This is a 
lignocellulose material in pellet form made by crushing apricot pits. 

Field-cleaning must be accomplished in accordance with published mainte- 
nance instructions for a specific engine. The first step is preparation of the 
engine for cleaning. On an installed engine, any coolers or instrumentation 
which may be damaged by the cleaning material must be removed from the 
aircraft inlet duct. Numerous other lines, tubes, sensing bulbs and controls 
must also be removed to prevent accumulation of cleaning material which 
might cause a malfunction in either aircraft or engine-mounted equipment 
or instrumentation. Airbleed ports and other openings should be covered. 
Usually, a 55-gallon steel drum is installed as a hopper above and slightly 
forward of the engine air inlet. Fwo hoses of proper size, fitted with simple, 
remotely operated valves, are connected from the bottom of the drum to the 
engine air inlet. The drum is filled with a specified amount (usually 100 
pounds) of Carboblast, and the engine is started. 

The engine is operated, first, at Idle, then, at higher rpm for stipulated periods 
of time. The total engine running time will probably be in the vicinity 



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of 15 to 20 minutes. After being cleaned, the engine is shut down to remove 
any caps or closures which were temporarily installed over bleed or pressure 
line openings. The engine is again operated for several minutes to blow out 
any cleaning material which might have accumulated in the bleed or pressure 
ports. The previously removed engine and aircraft equipment is reinstalled, 
following which the engine is ready for service. It should be noted that dual 
axial compressor engines must be retrimmed after being cleaned. 

RPM for Checking Thrust on Dual Axial 
Compressor Engines 

Even though turbine discharge pressure or engine pressure ratio is used as 
the principal engine operating variable for checking or setting the thrust 
produced by a dual axial compressor turbojet engine, an inoperative pres- 
sure gage may occasionally necessitate that rpm be substituted for this pur- 
pose. Except for the complications involved, the inevitable inaccuracies and 
the ever-present chance of error in the required computations, there is no ob- 
jection to this method, provided the engine operator understands the procedure. 

The Temperature-Rpm (or speed bias) Curve (Figure 4-3) may be used to 
determine the per cent of engine trim speed (based upon sea level static 
standard day conditions) at which various amounts of engine thrust, such 
as Normal Rated or 80% Normal Rated, will be produced for the exist- 
ing ambient (outside) air temperature or compressor inlet total temperature. 
However, the per cent of trim speed under standard conditions for a given 
engine thrust and inlet temperature, as computed from the curve, cannot be 
used directly for setting the throttle or determining the engine thrust being 
developed, To do so would be to assume that the per cent of trim speed under 
standard conditions, determined from the curve, would be the same as the 
aircraft tachometer reading. However, the trim speed will vary among en- 
gines and, additionally, will not be the 100 per cent reading on the tachom- 
eter. The value obtained from the curve, then, must be corrected so that it may 
be applied to the aircraft tachometer in terms of per cent of engine thrust. 
The actual trim speed of the individual engine, obtained from ‘the engine 
data plate, is used to make this correction. 

As a matter of good practice, the “field trim speed” obtained from trimming 
an installed engine in the field should not be used for computations when 
rpm is used as a measure of engine thrust. There is the ever-present chance 
that the latest “field trim speed” may not be accurately known to the pilot, 
or that the compressor may have been “field-cleaned” and trimmed without 
the pilot’s knowledge. It is therefore recommended that the data plate trim 
speed and not the “field trim speed” be used for correcting the standard day 
sea level static per cent trim speed obtained from the Temperature-Rpm 
Curve. Although the use of data plate trim speed, rather than the latest “field 
trim speed” may result in lower throttle settings than might otherwise be 
permissible, the error will always be on the “safe” side and will not endanger 
the useful life of the engine. 

It’s another story if an uninstalled engine is being calibrated in an engine 
ground test stand. Here, the greatest possible accuracy should be the rule. 





In this case, the latest actual field trim speed in rpm is used in conjunction 
with the Temperature-Rpm Curve. This must be converted to per cent to be 
compatible with the curve and the tachometer reading. Divide the trim speed 
in rpm by the rpm value at which the tachometer reads 100 per cent, obtained 
from the applicable P&WA Specific Operating Instructions or aircraft Flight 
Handbook. Using this per cent, correct the per cent standard day static trim 
speed determined from the curve. 

To use the Temperature-Rpm Curve (Figure 4-3), enter the curve with the 
ambient (outside) air temperature or, if airborne, with the free air tempera- 
ture gage reading, correcting, if necessary, to include the full temperature 
rise due to ram, to obtain the compressor inlet total temperature (T,.,). Pro- 
ceed vertically to intersect the curve representing the desired or predeter- 
mined required throttle position. This might be, for instance, Military, or 
Take-off (dry) for use during take-off, or possibly 80% Normal Rated 
for a cruise thrust setting. From the intersection of the ambient temperature 
projection with the desired throttle position curve, move horizontally to the 
left to read the per cent of trim speed under sea level static standard day 
conditions for the given ambient or free air total temperature. Obtain the en- 
gine trim speed in per cent from the data plate. Multiply the per cent of 
trim speed just computed by the per cent trim speed from the engine data 
plate to obtain the tachometer reading for the desired throttle position and 
prevailing compressor inlet total temperature (T,.,). The latest field trim 
speed in per cent may be used, as explained, when necessary for accuracy in 
a test stand. Adjust the throttle to obtain this corrected reading on the air- 
craft tachometer. 

Exhaust Gas Temperature Sense 

Experience has shown that a definite relationship exists between excessive 
exhaust gas temperatures and premature engine removals. The fuel control is 
designed in such a manner that, normally, exhaust gas temperatures will be 
maintained within a safe margin. However, the control cannot compensate 
for operational malpractices. Furthermore, under extreme flight conditions 
or in the event of a malfunction, the regulation of engine internal tempera- 
tures can be marginal or even above the desired limits. It behooves the pilot 
to develop a “temperature-sense” by thinking in terms of the interrelation- 
ship of the engine control factors. By so doing, he can soon learn either to 
avoid excessive temperatures altogether or to take immediate corrective 
action should they occur. 

It is sheer folly to treat overtemperaturing lightly. Just because the turbine 
does not fly apart or the engine melt away, there is no reason to assume that 
the engine cannot be or has not been damaged (Figure 4-5). Several momen- 
tarily high overtemperatures will have as profound an effect on the engine 
as a single prolonged one of a lesser degree. Excessive internal temperatures 
aggravate such conditions as creep, deformation of sheet metal parts and 
drooping. Operating the engine within the specified limits of temperature, 





_ Em Pea DP we Eee Bae Se BS SP See Se SS Se SS 

—_— TZ Tw» 





Rubbing caused by 
elongation or creep 

rpm and turbine discharge pressure or en- 
gine pressure ratio should become an in- 
stinctive technique to the turbojet or turbo- 
prop pilot. 

On some engines, exhaust gas temperature 
will increase with altitude, for a given 
throttle position. Retarding the throttle will 
usually maintain the temperature within al- 
lowable limits whenever a tendency to over- 
temperature is encountered. It may be pos- 
sible to control excessive acceleration tem- 
peratures by means of a slower throttle 
movement. The Specific Engine Operating 
Instructions published by the engine manu- 
facturer, the aircraft Flight Handbook or 
the applicable military Technical Order will 
stipulate allowable exhaust gas temperature 
limits for various operating conditions. In 
many instances, such as for take-off, opera- 
tion at Military Rated thrust or during an 
acceleration, the allowable maximum tem- 
peratures are time-limited. Whenever allow- a. 

able temperatures cannot be maintained or Turhine Blade mameye Cause 
controlled, the engine should either be shut by Excessive samparayres 
down or a landing should be made as soon as possible. Overboard tempera- 
tures must be reported by the pilot as a flight discrepancy, it being particu- 
larly important to record the peak temperature reached and the length of 
time that the temperature was above the allowable maximum, in order that 
prescribed maintenance inspections may be made after landing. 

caused by creep 

A pilot’s best assurance that his engine will render long and dependable 
service is to maintain engine operation always within the limits of tempera- 
ture and rpm. Whenever possible, it behooves him not to use a thrust or 
power setting higher than is necessary to accomplish his assigned mission 
because engine life is closely associated with temperature and rpm, even 
when the allowable limits are not exceeded. The “time-temperature-rpm” re- 
lationship within the engine is an important factor in engine durability. The 
most important of these is temperature, although all three enter the picture. 

The strengths of the materials used in the engine decrease as high internal 
temperatures climb toward the melting points of the metals, even though the 
danger point may not be approached closely. All materials stretch when a 
tensile or pulling load is applied. They also bend. when loaded as a beam. 
There is a tendency for any material to take a permanent set, stretch or bend, 
the tendency increasing with both the load and the temperature. The amount 
of permanent set increases with the length of time that the load and/or tem- 
perature is applied. After a certain amount of permanent set is attained, the 
fibers or grains of the material begin to pull apart. Under inspection in a 
highly powered microscope, the beginnings of fine cracks may be seen. With 


Blade angle flattened 

Necked down area 

Intergranular cracks 
are first indication 
of imminent failure 




additional time, the material begins to elongate at faster rates as the cracks 
become bigger and deeper. Finally, the material breaks. This process is so 
slow that elongation is perceptible only with careful measurement. The term, 
“creep,” has been applied to the process because of the length of time re- 
quired for elongation to become significant. 

In a turbine engine, high load and high temperature are usually experienced 
at the same time. The loading on the turbine and compressor blades is 
principally the combined result of the centrifugal force, associated with rpm, 
and some gas or air load, associated with engine internal pressure. When the 
turbine discharge pressure, which is indicative of other internal pressures, is 
high, so also will be the exhaust gas temperature. This means that when the 
turbine blades are subjected to their heaviest load, the material of which they 
are constructed will be its weakest. The compound effect of high rpm and 
high temperature results in an astounding increase in the rate of creep at very 
high thrust or power settings when the centrifugal load is the greatest. The 
ends of the compressor blades and the rims of the turbine wheels tend to 
travel outward. The rate of creep increases tremendously as the rpm and the 
exhaust gas temperature approach maximum. Numbers can be assigned to 
the relative amounts of creep to show what actually happens at varying ex- 
haust gas temperatures and engine speeds. For a typical turbojet engine tur- 
bine blade, the rate of creep is approximately as follows: 

Maximum Allowable 

1 unit per hour 

5 units per hour 
50 units per hour 
2500 units per hour 

Turbine life is directly proportional to the number of creep units per hour. 
The pilot controls the magnitude of creep by the manner in which he op- 
erates the engine. Turbine blades are carefully inspected and measured at 
engine overhaul. Those which have elongated beyond tolerable limits and 
those which show evidence of distortion or cracks must be replaced. In ex- 
treme cases, the blades may even fail before the engine becomes due for 
overhaul. It can be readily seen from the above table that when an engine is 
operated at the lowest temperature and rpm shown, the turbine blades will 
last 2,500 times as long as they will if the engine is operated at the highest 
temperature and rpm shown. Although an engine will operate satisfactorily 
at the maximum allowable temperature and rpm, it is obvious that the 
operating time between turbine blade replacement and other possible engine 
difficulties will be greatly increased if conservative engine operation is the 
rule and not the exception. 

Compressor Stall 

Compressor stall (or surge) has already been discussed in some detail. There 
are many different types of stall and their theoretical aspects are so compli- 




> Ga 

—_ Ge 



cated that no general division of type nor hard and fast rule pertaining to 
their cause and method of handling the engine when in a stall condition can 
be made. Sometimes a stall may be an indication of an engine malfunction. 
At other times, stall may occur by itself and may be nothing to be overly 
concerned about. Because most pilots may sooner or later experience com- 
pressor stall, a pilot should know what he can do to eliminate the condition. 
Compressor stall will usually be associated with unstable engine operation 
and may be recognized by compressor pulsations felt through the aircraft 
structure. In some instances, stall may even cause very loud explosive “bangs” 
to be emitted from the engine, which, to a pilot experiencing a severe engine 
stall for the first time, will be startling, to say the least. This is especially true 
in a single-engine aircraft. In a multiengine aircraft, a stall in only one engine 
will not be as noticeable. 

As disconcerting as a stall may be to the pilot at the time, however, there is 
no evidence to date that a Pratt & Whitney Aircraft turbojet engine has ever 
suffered damage as a direct result of compressor stall. The high thrust output 
and low specific fuel consumption of most axial compressor engines neces- 
sitate that the engine operate as close as possible to the stall region of the 
compressor. Designing close to the stall margin means that stall may occur 
occasionally in spite of all of the precautions that the design engineer has 
taken to eliminate the possibility. The amount and frequency of stall are de- 
pendent upon a number of things. Except in the case of an “Off-Idle” stall, 
low inlet air temperatures and the effect of Reynolds number at high altitude 
increase the tendency of an engine to stall. Consequently, stall recovery is 
better at low altitudes. Under some flight conditions, such as when the airflow 
pattern to the engine air inlet is at a high angle of attack, when, for instance, 
the aircraft is slipping or skidding or in the event of some obstruction like ice 
or foreign matter in the duct, the flow of air through part of the compressor 
slows down, resulting in a stall. Quite frequently on some engine installations, 
a mild transient stall may be experienced when the engine is accelerated from 
Idle to the thrust range just above Idle. Unless the engine accelerates 
too slowly, this so-called “choo-choo” or Off-Idle acceleration stall is of 
no serious consequence. 

A compressor stall will usually be accompanied by a loss of thrust which 
is reflected by the aircraft instruments, by a rapid reduction or fluctuation 
of rpm at a constant throttle position or failure of rpm to continue to in- 
crease during an acceleration, and by a rise in exhaust gas temperature. No 
response in turbine discharge pressure or engine pressure ratio to an increase 
in throttle position is another indication of compressor stall. 

There are several things that the pilot can do to avoid stall or to reduce its 
intensity. Erratic and abrupt throttle movements should be avoided. Rapid 
throttle advances during periods of high distortion of the air entering the 
air inlet duct, such as at low airspeeds, are sometimes the cause of accelera- 
tion stalls. Carefully coordinated flying increases the efficiency of the com- 
pressor inlet air duct. Airspeeds should be maintained above the acceptable 
minimum. Once a stall occurs, slowly retarding the throttle to Idle and then 
slowly advancing it to obtain the desired thrust may correct the condition. 





This is particularly true in the case of an Off-Idle stall. Operation at a 
higher airspeed and at a reduced angle of attack or rate of climb may elimi- 
nate the stall. As experience is gained and a “feel” of the engine is acquired, 
an instinctive sense of how to handle the engine in order to avoid possible 
stall-producing conditions becomes almost second nature. If the stall or 
unstable engine condition cannot be controlled, either the engine should be 
shut down or a landing should be made as soon as possible because con- 
tinued severe compressor stall could conceivably be detrimental to the engine. 

Turboprop Characteristics 

In many respects, a turboprop may be considered as just another gas tur- 
bine engine. Operationally, however, this type of engine has some funda- 
mental characteristics which make it quite different from the standpoint of 
the pilot. 

A turboprop engine combines the advantages of a turbojet engine with the 
propulsive efficiency of a propeller. The turbojet engine derives its thrust by 
rapid acceleration of a relatively small mass of air. The turboprop develops 
propulsive force by imparting less acceleration to a relatively large mass of 
air. The turbine of a turbojet engine extracts only the necessary shaft horse- 
power to drive the compressor and the accessories. The turbine of a turbo- 
prop is designed to absorb large amounts of energy from the expanding 
combustion gases in order to provide not only the power required to satisfy 
the compressor and other components of the engine but to deliver the maxi- 
mum torque possible to a propeller shaft as well. Propulsion. is produced 
through the combined action of a propeller at the front and the thrust pro- 
duced by the unbalanced forces created within the engine that result in the 
discharge of high-velocity gases through a nozzle at the rear. The propeller 
is responsible for roughly 90 per cent of the total thrust under sea level static 
conditions on a standard day. This percentage varies with airspeed, exhaust 
nozzle area and, to a lesser extent, temperature, barometric pressure and 
the power rating of the engine. The power supplied to the propeller is meas- 
ured as shaft horsepower (shp), to which must be added the effect of jet 
thrust when the total power output or equivalent shaft horsepower (eshp) of a 
turboprop engine is calculated. Performance calculations for turboprop engines 
have been discussed and illustrated in Section III, to which reference should 
be made for information pertaining to the calculation of turboprop horsepower. 

Although some turboprop engines employ a compressor of the centrifugal 
type, larger, high-performance models almost invariably require the greater 
efficiency and higher compression ratios attainable only with an axial-flow 
compressor. The compressor may be either of single or dual rotor design, 
the latter having both a low pressure compressor and a high pressure com- 
pressor. When a single compressor is used, the propeller reduction and drive 
gear is usually connected directly to the compressor shaft, and, when a dual, 
or so-called split, compressor is used, it is connected to the low pressure rotor. 
Sometimes, the propeller is driven independently of the compressor by a free 
turbine of its own. 






In spite of the fact that it is more complicated and heavier than a turbojet 
engine of equivalent power, the turboprop will deliver more thrust up to high 
subsonic speeds (Figure 4-6). The advantage decreases as airspeed increases. 
In normal cruising speed ranges, 
the propulsive efficiency of a 
turboprop remains more or less 
constant, whereas the propulsive 
efficiency of a turbojet increases 
rapidly as airspeed increases. 
The spectacular performance of 
the turboprop during take-off 
and climb is the result of the 
ability of the propeller to accel- 
erate a large mass of air at rel- 




f atively low flight speed. 
| : AIRSPEED-KNOTS s00 Jf it is assumed that the fuel 
(ENGINES OF THE SAME APPROXIMATE SIZE) flow of two ae turbine engines 
| Figure 4—6. Comparative Net Thrust at of the same size will be sub- 
Sea Level stantially the same under simi- 

lar conditions, it follows that the 
one equipped with a propeller and consequently delivering the most thrust 
will have the lower thrust specific fuel consumption (tsfc). In the 400- to 
500-knot speed range, this supremacy can be as much as 20 per cent or more 
(Figure 4-7). The turboprop at- 
tains its most economical oper- 
ation at a somewhat lower air- 
speed than a turbojet of equiv- 
alent power. Turbine power- 
plants develop their best power 
and efficiencies at high rpm. In 
the case of the turboprop, full 
power for take-off is obtained 
at compressor speeds four times 
that of a reciprocating engine 
crankshaft under similar condi- 




me TS Ge 

400 650 

| efficiency iS produced only when (ENGINES OF THE SAME APPROXIMATE SIZE ) 
the engine is operating within a Figure 4—7. Comparative Thrust Specific 
narrow range of high rpm. The Fuel Consumption 

normal operating range of var- 

ious types of engines differs greatly. The reciprocating engine develops its 
propulsive power gradually over a relatively broad range of rpm as com- 
pared with that of the turboprop (Figure 4-8). This is because the efficiency 
of a vane-type compressor, whether centrifugal or axial, is dependent upon 
high rpm, whereas the efficiency of a piston-type compressor is relatively 
insensitive to the speed of piston travel. Progressively larger amounts of 
turboprop power are obtained by increasing the propeller blade angle and 
fuel flow ratlier than by increasing rpm. 

vik Ee ah 






INSTR. 200 

Figure 4-9, portraying typi- 
cal turboprop performance 
as it relates to throttle set- 
ting, gives an indication of 
what to expect when this 
type of engine is operated. 
The curves, as presented, are 
more or less representative 
of the characteristics of a 
Pratt & Whitney Aircraft 
Figure 4—8. Reciprocating and Turboprop P12 or T34 turboprop en- 
Engine Operating Ranges gine. The curves will change 
somewhat when different 
fuel controls are used. Typical curves for other turboprop engines (the rpm 
curve, in particular) will not be the same although they, in all probability, 
will reflect the same general engine characteristics. 

( Turboprop ) 


The features and operation of 
a turboprop fuel control and 
propeller governor have been 
described. The flight operating 
and ground (Beta) -operating 
ranges of the fuel control are 
sometimes called the propeller- 
governing and propeller-nongov- 
erning ranges, respectively. The 
operation of the fuel control 
and the propeller governor 
should be understood by the pi- 
lot if he is to obtain the most 
from his engine. 


Engine RPM- % 



Although the following discus- 
sion pertains generally to turbo- 
prop engines of all types, the 
characteristics described are 
mostly those of the Pratt & 
Whitney Aircraft PT2 (or T34, 
as the military version is known). . 
This engine is considered more 

or less typical of relatively large, Figure 4—9. Typical PRWA PT2 or 734 
single-axial-compressor, directly _ Turboprop Performance 
coupled turboprop engines. 

*——— Reverse —** Forward 


Four definite throttle positions are normally provided, with a quadrant detent 
for each. When the throttle is fully forward in the Take-off position, the 
engine will produce maximum power. This is usually a time-limited rating. 
“Flight Idle” is the minimum setting permissible for airborne operation. 
“Ground Idle” represents the best rpm for normal ground operation. “Re- 



Shaft Horsepower - % 


| “= "2 ma wma EZ Ge a | | “= “= = Ge _= a Gwe 




verse” is what the name implies; that is, the throttle position at which maxi- 
mum reverse thrust will be obtained by means of reversing the propeller blade 
angle. Setting the throttle at any desired intermediate position on the quadrant 
is permissible. Once the throttle has been set in the position required, the fuel 
control and the propeller governor will establish and maintain the power 
called for by the pilot. 

Both Ground Idle and Flight Idle thrust and rpm values may vary greatly 
with different types of fuel controls and with different engine installations. 
Thrust and rpm values are preset in the fuel control and propeller governor 
to match the characteristics of a particular airplane as well as to produce 
predetermined performance. 

The Turboprop Propeller 

The propeller of a turboprop engine retains most of the essential features 
common to those employed on large piston engine installations. Both hydro- 
mechanical and electrically controlled propellers are in current use. From an 
operational standpoint, the differences between the two types are minor. As 
has been mentioned, in a turboprop engine, the fuel control and the pro- 
peller governor are coordinated. The propeller and engine rpm are mechani- 
cally governed in the flight operating range. In the Beta or ground operating 
range, propeller pitch varies with throttle position. Propeller blade angles 
from full feather to full reverse pitch may be obtained throughout the entire 
range of operating rpm. Because of the high rpm of a gas turbine engine, a 
reduction gear arrangement, similar to that described in Section II, is usu- 
ally used. 

The blades of a turboprop propeller must have very rapid pitch-changing 
characteristics. The narrow rpm operating range (Figure 4-8) of the engine 
requires that the propeller blades change angle at a much more rapid rate 
than is required in the case of the reciprocating engine. The blades of a 
reciprocating engine propeller have approximately 40 per cent of the rpm 
range in which to vary their pitch from 20° to 45°, while the blades of a 
turboprop propeller must change from about 5° to 45° in only 10 per cent 
of the rpm range (Figure 4-10). Translated into flight operating technique, 


20° =Operational Low 
Blade Angle Stop 
a ROTATION (optional 
Idle —5° Flight Idle 
(Reciprocating) ( Turboprop) 

Figure 4—10. Propeller Blade Angle Variation 

this means that the turboprop engine is much more sensitive to throttle 
movement than a reciprocating engine. 



ee ee, ee ee ere 



The turboprop propeller blade angle at Flight Idle is small when compared 
with the blade angle setting (approximately 20°) for a reciprocating engine 
propeller during a glide at minimum power. The turboprop airplane con- 
sequently can have high aerodynamic drag, provided that the fuel control 
and propeller governor are adjusted to provide this characteristic during 
glide and approach. High drag will result in a faster rate of descent than 
might be attained by a similar airplane powered with reciprocating engines. 

Unlike a reciprocating engine, a turboprop has no propeller control lever, 
as such. However, a three-position lever or switch is customarily provided 
for feathering, unfeathering and normal (automatic) operation of the pro- 
peller. This is usually known as the propeller condition lever (or switch). 
When in the unfeather, or what is frequently called the airstart, position, the 
propeller moves to a preselected blade angle suitable for an in-flight engine 
airstart at any combination of airspeed and altitude. 

Turboprop Operation 

Turboprop engine operation is quite similar to that of a turbojet engine 
except for the added feature of a propeller. The engine preflight inspection, 
the starting procedure and the various operational features are very much 
alike. The turboprop in flight chiefly requires attention to engine operating 
limits, the throttle setting and the torquemeter pressure gage. Although 
torquemeters indicate only the power being supplied to the propeller and 
not the equivalent shaft horsepower (eshp), torquemeter pressure is approxi- 
mately proportional to the total power output and hence is used as a meas- 
ure of engine performance. The torquemeter pressure gage reading during 
the take-off engine check is an important value. It is usually necessary to 
make a preflight take-off power computation in the same manner as is done 
for a dual axial compressor turbojet engine. This computation is to determine 
the maximum allowable exhaust gas temperature and the torquemeter pres- 
sure which a normally functioning engine should produce for the outside 
(ambient) air temperature and 
barometric pressure prevailing 
at the time of take-off. Take- 
off power curves, similar to 
Figure 4-11, are used to make 
the computation, as shown by 
the example. On a cold day, if 
a torque limiter is not a feature 
of the fuel control, it may be 
found that the take-off must be 
made at a throttle setting below 
that of the normal take-off posi- 
tion on the throttle quadrant in : 
order not to exceed the allow- 

able torquemeter pressure for TAKE-OFF POWER CURVE 
the prevailing conditions. When Figure 4—11. Turboprop Take-off 
these or similar curves are used, Power Curve 


ou 4. 


Y < 

| a Hm PP mH EH a fm aD HD HR aD an aD an Gan GH ca= Guan —— aa 

—_ ae Mm Te Dae De ae Pw we BP ee DP ewe Se Se S&S 





it must be remembered that field barometric pressure and not the barometric 
pressure corrected to sea level (as normally given by airport control towers or 
air weather stations) must be used. The horsepower available for take-off 
may be computed by employing a torquemeter constant and multiplying by 
the actual rpm (not per cent) and by the torquemeter pressure reading either 
in inches of mercury or pounds per square inch (as the case may be). 

Climb and level flight are accomplished at appropriate throttle settings. 
Cruise power settings for various altitudes and desired airspeeds are obtained 
for known gross weights by reference to charts or curves furnished by the 
aircraft manufacturer. The most favorable altitude for any given weight 
depends upon the ambient air temperature. Consideration must be given 
to the free air total temperature, the pressure-altitude and the airspeed. The 
required torquemeter pressure for any given set of conditions is then deter- 
mined from cruise control charts, and the throttle is adjusted accordingly. 
Although these procedures may, at first, seem complicated, the task is not 
as difficult as it might appear, and once familiarity has been acquired with 
the process, there should be no problem in computing operating data for the 
engine under any flight condition. 

It has been mentioned that it is essential that the propeller be maintained 
in the propeller-governing range while the aircraft is airborne and that the 
Flight Idle position is the lowest throttle setting before the propeller enters 
the nongoverning or Beta regime. The Flight Idle position, then, is the mini- 
mum setting permissible for glide conditions. A throttle stop is often pro- 
vided to prevent the throttle from being moved below the Flight Idle position 
until after the aircraft has landed. The rpm at Flight Idle is sufficiently high 
to ensure that the engine is maintained in a constant state of readiness to 
respond effectively to a demand for full power should a “so-around” become 
necessary during a landing approach. 

The fuel control sometimes may be so adjusted that the fuel flow to the 
burners, and, consequently, the power supplied by the turbine will not be 
sufficient to supply all of the power required by the compressor at minimum 
airspeed for glide. The balance of the power required must be absorbed by 
the propeller from the forward motion of the airplane. In other words, when a 
glide is accomplished at Flight Idle, a negative thrust will be produced which 
will increase the normal glide angle. A flatter angle of glide or approach, 
or a lesser rate of descent or sink, is easily obtained by using more power. 

In a turboprop-powered airplane, the decision as to whether or not reverse 
thrust will be needed after landing should be made while on the final ap- 
proach. Immediately following touchdown, the throttle is moved either to 
the Ground Idle position or to the desired reverse thrust position, either full 
or partial reverse being available. Reverse thrust is most effective at high 
aircraft speed. As forward velocity is lost, the effect of reverse thrust deterio- 
rates quickly. It should be borne in mind that airspeed, not ground speed, 
determines this effectiveness. When an approach is made into a high wind, 
large amounts of reverse thrust relative to the initial ground speed may be 





anticipated. The opposite would be true in the event of a down-wind landing. 
Low pitch stops, usually incorporated in the propeller, will maintain a blade 
angle sufficient to ensure a large amount of aerodynamic braking at touch- 
down even though reverse thrust is not employed. At an airspeed of 100 
knots and with the throttle in the Ground Idle position, this effect may pos- 
sibly be as high as 60 per cent of full aerodynamic reverse thrust. The effect 
falls off rapidly with diminishing airspeed. 


Gas turbine engines are normally not prone to complete engine failure. In 
fact, they have the fortunate characteristic of continuing to produce high 
thrust or power even when suffering from a malfunction which, if compara- 
ble in severity, might have long since caused a reciprocating engine to fail. 
Literally millions of flying hours have conclusively proved that engines of 
this type are remarkably reliable. Yet, malfunctions do sometimes occur 
and it behooves the pilot to remain alert for possible trouble even though 
the chances are remote. 

In the event of simple engine flameout, gas turbine engines may be consist- 
ently relighted in flight at altitudes up to 40,000 feet, and above. Internal 
engines fires in flight are rare. Should one occur during engine ground opera- 
tion, the accepted practice is to shut off all fuel, as well as the ignition, and 
then blow the fire out of the tailpipe with air from the compressor by turning 
the engine over with the starter. The appropriate Flight Handbook, engine 
Operating Instructions or the applicable military Technical Order should be 
consulted for emergency procedures and techniques. 

The maintenance of gas turbine engines should never be.of the trial-and-error 
variety. Since the internal operation of this type of engine is based upon 
involved thermodynamic and aerodynamic theory, it will not do to trouble- 
shoot by means of the hit-and-miss tactics of tearing out or adjusting sus- 
pected parts or systems at random. A malfunction must be carefully analyzed 
and its cause reasonably ascertained before the engine, itself, is attacked 
with tools. It therefore behooves the pilot and his flight crew to report the 






) ) 


ee ee es 



symptoms of a malfunction carefully, including an accurate recording of the 
instrument readings and other pertinent data both before and immediately 
after the incident in order to provide maintenance personnel with sufficient 
information to determine the cause of the trouble intelligently. 


When the development of gas turbine engines was in its infancy, it was gen- 
erally believed that the engines could operate on almost any kind of fuel. 
Statements were made to the effect that nearly anything from crude oil to 
aviation gasoline of 115/145 grade could be used successfully. As experience 
was gained, however, it soon 
became obvious that specifi- 
cations, closely controlling 
important qualities of the 
fuel, would be necessary if 
maximum performance was 
to be obtained from engines 
of the gas turbine type. 
Through the years, the speci- 
fications have become pro- 
gressively more specific. 

At first, aircraft gas turbines 
almost exclusively used what 
has been loosely described 
as kerosene. In the United 
States, the term, kerosene, is employed to describe a very wide range of 
petroleum defined only by a minimum flash point of 120°F and an end 
point of not more than 572°F. Flash point is the fuel temperature at which 
sufficient vapor forms at the surface of the liquid for the vapor to ignite in 

. JNSTR. 200 

air when a flame is applied. End point is the fuel temperature at_which all 
of the liquid will distill over into vapor. There are specifications for kerosene- 
type fuels for certain applications, such as lamps or heaters, which are some- 
what more explicit than the simple requirements relative to flash point and 
end point. Nonetheless, the term, kerosene, does not closely define a fuel. In 
general, the kerosene available in Europe and other parts of the world is 
cleaner and more suitable as a gas turbine engine fuel than the average Amer- 
ican kerosene has been. When proper requirements are stipulated in the fuel 
specification, high-quality kerosene can be produced in this country. Today’s 
gas turbine fuel specifications define both kerosene and other high-energy 
fuels suitable for aircraft engine use. Since aircraft gas turbines are designed 
for operation on specific fuels, only the fuel specified by the engine manu- 
facturer, the aircraft Flight Handbook or the appropriate military Technical 
Order should be used. 

Military Fuels 

The fuels used in gas turbine engines by the military services have been given 
the prefix, “JP.” A brief review of these fuels will illustrate the evolution that 





has taken place in the development of aircraft gas turbine fuels since the 
engines were first introduced. The JP-1 and JP-2 fuels are no longer avail- 
able as such. 

JP-1 was a kerosene-type of fuel of low freezing point. Because the low 
freezing point (—76°F) requirement made procurement difficult, produc- 
tion of this fuel was rather limited. 

JP-2 was somewhat more volatile than JP-1, but was never used extensively. 
The specification was written late in World War II as a means of relieving 
a potential shortage of JP-1. 

JP-3 was developed as a fuel of maximum availability and was, as a result, 
more volatile than JP-2. Because of its volatility, great losses were experi- 
enced in flight through evaporation at high altitude and during high rates 
of climb. 

The specification for JP-4 was issued in 1951. The fuel has a desirable 
lower volatility than JP-3 but was, at first, a step backward toward JP-2 
because the JP-4 specification was more restrictive as far as availability was 
concerned. Later, some of the properties of the fuel other than volatility 
were altered to ease the supply situation. 

JP-5 was developed as a heavy kerosene to be used as a product to be 
blended with gasoline to produce a fuel similar to JP-4. This procedure 
was desirable for naval operation aboard aircraft carriers where only limited 
storage space was available for the necessarily large quantities of JP-4. The 
JP-5 fuel could be, blended, as needed, with the supply of aviation gasoline 
carried aboard ship for aircraft powered with reciprocating engines. The 
JP-5 could be stored in any available tanks normally used for the ship’s 
diesel engine fuel. In emergencies, it was also possible to use the JP-5 in 
the ship’s own engines. However, as aircraft gas turbine development pro- 
gressed, even JP-4 became too volatile for some operations, so engines were 
designed to use JP-5 directly. The JP-5 has a high flash point (140°F) 
and very low volatility characteristics. 

JP-6 was developed by the Air Force for land-based, high-speed supersonic 
aircraft. The fuel is slightly more volatile than JP-5 and has a low freezing 
point (—65°F) for aircraft operation in cold climates and at high altitudes. 

Commercial Fuels 

When it was assured in 1956 that gas turbine engines would be used in 
commercial aircraft, it became evident that a definition was needed for a 
fuel suitable for commercial airline operation. At the present writing, the 
American Society for Testing Materials (A.S.T.M.) has published, for 
informative use only, specifications defining two grades of fuel suitable for 
commercial use. One is a JP-4 type of fuel and the other is a JP-5 or kero- 
sene-type fuel. Pratt & Whitney Aircraft has issued a specification, designated 
PWA 522, which defines the properties needed in the fuel for the dual axial 
compressor engines which will power commercial turbojet aircraft. The re- 
quirements of the PWA 522 specification, together with any additional re- 


Ss u 


3 J 5 5 5 3 

— Ge 

—_= a aS we 


PWA OPER. | NSTR. 200 

quirements desired by individual airlines, will provide an adequate specifica- 
tion for various commercial airline operators to employ when purchasing gas 
turbine fuel. In all likelihood, each operator will use a slightly different pur- 
chase specification to suit the needs of his particular operation. Commercial 
gas turbine engine fuel will therefore embody a range of volatility character- 
istics similar to that encompassed by both the JP-4 and JP-5 military fuels. 


The vast background of experience that has been gained in the field of air- 
craft engine lubricants as a result of years of reciprocating engine operation 
is confined to noncompounded mineral oils. In general, these oils are very 
viscous, similar in viscosity to SAE 50 and 60 grades. During cold weather, 
reciprocating engine starts can be accomplished without preheating the engine 
only when the oil has been diluted with fuel at the time that the engine was 
last shut down. The gasoline used to dilute the oil subsequently evaporates 
when the oil again becomes heated. Hence, the change in the characteristics 
of the lubricant is only temporary. 

Gas turbine engine fuels are 
not sufficiently volatile to be 
safely used for oil-dilution 
purposes. Therefore, a new 
lubricant had to be found 
which would flow at tem- 
peratures as low as —65°F 
and still lubricate satisfac- 
torily at temperatures con- 
siderably higher than those 
experienced in reciprocating engines. The major advantage of the synthetic 
oil which has been developed to fill this difficult requirement is that it is an 
effective lubricant over a much wider temperature range than its mineral 
oil predecessors. 

Synthetic oil for gas turbine engines can be manufactured from any of 
several different basic materials, such as vegetable oil, petroleum oil and 
animal fats. Since some of these materials are not compatible with one 
another and since synthetic oils of different manufacture are not necessarily 
derived from the same base even though they meet the same specification, 
it is important to ensure. that synthetic oils produced by different manufac- 
turers are not mixed or indiscriminately used together in the same engine. 




. ij - t, LA ns 
. 7 oa v 7 
A = — a 
7 = pm 
i . - ¥ 
ra = 
. | ae ; ron 
. : j = 7 - 7 ¥ 7 WM = 
; ' a al gt = 2 , > | U 7 i 
» = . e062 2 7 = (3b teri U = ms 
: _ 4 as ‘ A i i = _ 7 fs - py © 2 
\ - | - j ” — - ae = a 
i is <P « a aA“ en ai O4 tio = 
: : 5 
Stner —y is to oe : ; 7 


Absolute temperature, definition ......... 74 
PC OUTIGT SIC reece 135 
Accessories, airbleed-driven. ........... 45 
Accessories, mechanically driven ...... 46 
Accessories, performance Of wc 103 
Adjustable exhaust nozzle cscs 43 
Aerodynamic thrust reverser ......sssssce 37 
Afterburner Check CUrVe ...cceescsssssssssssesees 103 
TIE ICE cccctinreecctercrriceretcien 42 
Afterburner fuel Control ......sscsessssessssseen 51 
Afterburner fuel system ....ccsssssssssscssnn 51 
Afterburner igniter ....sssssssssssssesssssssssseessees 52 
Afterburner Mozzle v.cccsccccccssssssessssssseeees 42, 43 
Afterburner nozzle Control .....sssm 43 
Afterburner performance ..ncccccsssssesessnsseees 102 
Afterburner SCree@ch ...sssssssssssssssseesssssssene 44 
eS a 12, 41 
Afterburning turbojet Engine ..........0« 12 
Air (free) temperature Gage... 116 
PT OU: FONE Ginn enteric 27 
Par NE NE isa enacntnornwiamione 16 
Air inlet duct performance ..rccsssssssm 101 
Si SN occ 59 
PE I i ccchnhiceneresieiinnnetinceinsicommetinn 142 
Airbleed, COMPLessor ..ncssssssssssssesenseseseeen 25 
Airbleed correction Factors ws 91 
Airbleed-driven Accessorie... 45 
Airbleed performance .neccecssssssssseseseeseeen 103 
PI NE aiaticice cn tctimenierecen 45 
Airflow Parameter occcccssssssssssseseseesnnsessssesee 90 
Alcohol injection .....ssssssssssesssessessnnsssssensseeee 66 
Altitude tables, stamdard ....ccccccssssssssessessees 104 
Analysis, turbojet engine perform- 
EI asic cies nisiencncinecipelapeiaianasinen 84 
Analysis, turboprop engine perform- 
OE a cicicncicnaiiceetidinaiaaneennn 92 
Annular combustion chamber ............. 28 
Annular-gap igniter... 63 
Anti-icing, engine inlet ......ssssssscsssseeee 121 
Anti-icing ene eae re 69 
Area adjustment, exhaust nozzle ..... 34 
Asymmetric thrust control, turboprop 70 
Axial COMPFeSSOLS o.nccccsssorsesssssssnsessenssseenne 21 


Basic gas turbine CMGiMe .....scccssssssn 4 
Bellmouth compressor inlet ....scsscsss 18 
Beta range, turboprop ...... 58, 138, 139 
Blade angle pitch stops, turboprop... 59 
Blade design, COMPLessOr ......ssssssssssn 22 
Blade design, turbine .....csssesssssussssssseen 32 
Blankets, imsUlation .ccccsssscccssssssssesecssssseeeeees 66 
NS RE cre cesiceeineenaoine 8] 
Brakes, turboprop CMGiMe .nccccvssscssssssuen 72 
ee i ee eer ae ee 60 
British thermal unit, definition ............... 74 
ee LL, Sea eer ae 32 
ee nee 29 
NS TI i ee 28, 29 
Burner performance Q..cseccssssssssseessesnssseseeeeee 102 
BUrNer SECTION, CMGIME oececcscccccssssssssseseeeeennes 27 
Burning time, combustion starter ........ 48 
i ON OO 14 
Calibrated lubrication system cs 59 
Calibration Curve, OMGiMe reece 125 
Can-annular combustion chamber ...... 29 
Seite DUIIOE cotiececretcenamnivtnecnennercnpen 28 
Cans, combustion (burner) eccssscsssssse 28 
Capacitor-type ignition system .......... 61 
Carboblast engine cleaning... 130 
Cascade effect, compressor blade ...... 22 
Case temperature, CNGiMe cscs 66 
Celsius (Centigrade) to Fahrenheit 
COPOIRIO Fini ecnctosmninencganaiaginn 105 
Celsius temperature SCONC ...ssssssseunsee 74 
Centigrade temperature scale .............. 74 
Centrifugal compressor-type engine 12 
Centrifugal COMpPressors  enacassssssssssssseen 20 
Chambers, COMBUSTION .......ssssssscscsscssssssseeeee 27 
CNN SU cc cictncicmepaciniecoeennns 81 
SAPS TINUE ccsvsiecticniteiectitnsincnes 52 
Clamshell thrust reverse .........csssssssssssssses 38 
Cleaning engine air passages ............ 130 
Climb, afte rburming ..ssescsssssscssssscssssessssees 119 
Climb and cruise (turbojet)... 118 
Closed cycle (CMgime) ...cscsseosssssnseseesnnsee 78 
Combustion Chambers. recscessssssesesee 27 
Combustion starters ...cccscssscssssssssssecesessessee 47 





Commercial engine ratings er. 111 
COPIRIUGIT TONING oasencscicscnssessessinssesinesniee 144 
Component performance details ......... 101 
Compression and expansion curves... 82 
Compression ratio, definition ccc... 75 
Compressor Girbleed .....ccesssssssesssssssessen 25 
Compressor inlet temperature vs true 
Compressor performance .eecescscsssssssnesen 101 
Compressor speeds, dual axial ............ 23 
Compressor stall] ....cscccsssssssssssssesenseee 25; 134 
Compressor stall COMO] ....ccssssscssssssssesen 55 
MII sail pchnsh cect cilities 19 
Condition lever, propeller (turbo- 
ID siiccinistcstiesesceucshecaiinaciaabababeini 140 
Conservation of Energy, Law of the 81 
Constant, turboprop torquemeter ...... 141 
Constraimed-gap igniter ...ccccsscccsssssssese 63 
Continvous flOW CY Cle oeccccssscsscsrsseneeensees 78 
Convergent-divergent exhaust nozzle 34 
Convergent exhaust Nozzle ....cccccsccs 34 
PCED TTI isssscncssisciccccscicctcmtenee 66 
OTe) ola) Mc -T° |] [>| 0) 69 
CNTR,  CINIINO oisniinciie cicieeces 64 
Creep, turbine blade .......... 32, 132, 134 
Cruise and climb (turbojet) 0... . 118 
Cruise, maximum (turbojet) cs... — oo 
Cruise power settings, turboprop ..... 141 
Cutoff speed, Starter ..rceeccssssscsssssssssessssseen 47 
NN, PIE ise casi ciceiaaieas 77 
Data plate speed, definition... 123 
Pecereiies, Gi so 60 
Delta, definition Of cecccccccccscssssssssesmeeen — 
ne NN? Fe 26 
Diffuser CASE IMJOCtiOM ceeesccoosseesssseseeeese 69 
Diffuser, centrifugal compressor ......... 20 
SN West chicedhkaciscnbiecmesicsepurdtionsits 26 
EOTPPOORNINGS ETN asc ssescsccinsescecssercniirece ae 

Discharge pressure, use of turbine .. 114 
Divergent exhaust nozzle (duct) ........ 35 

Divided entrance inlet GUE ...ecccccsssooue 17 
Double-entry COMPFeSsOr ...ccsccsssssseseneen 20 
Drain, fuel Mmamifold icccccccccsccssssssssseseeeees 50 
Dual axial compressor Engine -...cce..... 12 
Dual axial COMpressors ...ccecsccssssssssesesee 22 
ee COE eo 3] 
Duct, afterburner .......ccssssssscccscccccscssssssecesseceeee 42 
6 A Renn ee Te ee . 
neh: GE San i. 33 

Duct loss correction FOCEOR .nassoosssssssesse: 92 
Ducted fam Om gime onnneeccsssssssssssssssssossnsecseeen 14 
Dynamic balance, turbine ....cccccccccsssseoe 32 
eee 46 
Electronic fuel COmtrols ..ccscccssmsssessen 55, 56 
NII cccscsincincipsiaitnnssinsiciniiiiabidi 142 
Emergency fuel systerm ocecccscnccssssssesnsee 50 
End point, Gefimition ccccssssnsuessnsneeeen 143 
Energy exchanges, laws governing .. 80 
Engine airflow parameter nncccccesssssessssen 90 
Engine analysis, turbojet perform- 
SED ciniivicsadiaccapndiesniiatematbindveadinicinlineasanie 84 
Engine analysis, turboprop perform- 
SI dsccsistsssdacevapll ides mpmetiplteiaalngsihis 92 
Engine brakes, turboprop ..ccecccscesssses 72 
Engine instrumentation ....seecccssccsssssseeeees 114 
Engine operating variable .........ccs. 112 
Engine pressure ratio, use of .. 112, 114 
Engine ratings (Gemeral) ......ccccssscsscces 110 
UWEEPOD SIUITOWUIG eccccsessinaseeecnsasecrnesessosinnnee 121 
Engine station designations ... 10, 84, 92 
Engine tri CULVES ooncecccsssesssssonssee 126, 127 
Engine trim Speed ...ccsssssesssssessuseeesseeree 123 
Erie trite eeeesssssssssssccsesccsscceennone 34, 123 
Equivalent shaft horsepower, turbo- 
SI decsdaimitieineiaiictlabilatenedenniliaiebaneiagsio 140 
Excessive fEMPerature .........ccsccssorsoseeeeee 132 
ST POUR iii cassclctinsiesaschcbahcsnciei 62 
a a 33 
Exhaust gas pressure measurement... 33 
Exhaust Gas temperature ...ccccoocsscsseessen 115 
Exhaust gas temperature sense ............ 132 
Exhaust gas temperature, turboprop 140 
Exhaust mozzle Control ..ccccscsscccssssssssssseee 43 
TF  .  , 38 
Expansion and compression curve ...... 82 
Fahrenheit temperature scale .............. 74 
Fahrenheit to Celsius (Centigrade) 
III sccoshieansndidiccnseripcuaonstniacyneatbons 105 
FRE PCRTUNON ccenecssscecsssescsaceesrsosecctomseens 130 
Field trim speed, US Of ecscsessssssssesn 131 
Fir-tree turbine blades ...cseesmssssssssssesn 32 
Flame holders, afterburner .......cccsss0« 42 
FIAMEOUt, SMM eececsssenssssssesssssssesessneennees 142 
Flame tubes, Crime ecececsssscsssssssssssssessen 61 
Flash point, Gefimition ..eccscccsssesssssssnesse 143 

Flight Idle, turboprop. ........ 59, 138, 141 




= GD 

—_ Te 

_ & Ga GT 

= Te 

a Gy 


Flight operating range, turbo- 

COE: icdscctsharniny steele ite 138, 139 
Fhicgtt cebicghats | ancccscscsccssccosstsconsocsennsnsetencee . 142 
Flight Spectrum, the -cecsecsssssssseenseenssesenseees 1 
Flow matching, COMPLessOr eeccssssssseen 23 
Free turbine, turbOprop ..csssssssssssssseeen 13 
Paiel ‘alt: FO6 icin 27 
Fuel control, afterburner ......csssssssssssssse 51 
Fuel control FUNCTIONS .....ssssssssssssssceeeeeeenee 57 
Fuel control schedule ......sescssssssssseescsessen 54 
Fuel controls, basic CMGiMe ...ccscssssssssseseees 53 
Fuel controls, Gemeral .....essssssssssesssseeeeenn 48 
Fuel controls, tUrbOProp nesses 58 
Fuel flow, afterburner ..........ssssssssssessssssee 42 
Perel We HID einstein mrnestscetssssnizeceties 115 
Pisel RaW 00 OF esiicicientiicsinin 120 
Fuel icing prevention nccssccsssssssssseesnseeseeeen 70 
Pana CN So isnt cranes 26 
POE IIS chasis inescitnineenintinepenge 26 
Pica: CNIS 5 scsccissicsserecessincctricntiecsesotec 59 
PAGE CORINEES clans eisieenecticicnscg terriers 50 
Fuel pump inlet pressure gage ......... 116 
Fuel system, afterburner cscs 51 
Fuel systems, basic engine .........0. 48, 49 
Fuels, cas turbime .nnnssssssscssrsocesssceesnsscsernsee 143 
Gas generator, definition ....rccceecsesesn 11 
Gas turbine performance ..ercecsseen 82 
Gear boxes, accessory rive assesses 46 
General operating instructions, use 

CF Eainsisncnaotictind iceaattamnenmmmeandtiinia 109 
Glide angle, turboprop aircraft ......... 141 
Gross thrust coefficient, turbojet ......... 90 
Gross thrust, Gefimition .........sssssssssseceseesees 9 
Gross thrust parameter ...cccscsscsssseeeneseen 90 
Ground cart, Starter .....ccssssssssssseesssssssees 47, 48 
Ground Idle, turboprop .srersesssssu 58, 138 
Heat cind temperature .nnacesssossesseeensssnsee 73 
High pressure COMPFeSSOF sccm 23 
Horsepower calculation, turboprop ... 141 
Horsepower, definition Of ..eescccsesasennee 7 
Horsepower (shaft) correction factors 

do a, 1) eee cnn eee 99 
Horsepower (shaft) correction for 

How the gas turbine operates ........... 5 
How the gas turbime Works ..rcccssssssccseu 8] 
Hydromechanical fuel controls ..... 55, 56 


WSRPRD: EOROCHOR issisenisnncectensecerseneeserersstion 69 
Igniter, After bUrMer ..nssssssssssesssnssseeeeeeenn 42, 52 
CE TINIE sensccinicicocnstionnlite 61, 63 
VQMition SyStOTS ..n.nneccsssoneccccsseseorssnnssessesssees 61 
Impellers, COMpPLeSSOL ..cssssssssseeeseesseneeeeen 20 
IMpiNnGeEMeENt StCTOLS ...ccorecseeeseeeceensseeeeens 46 
Impulse Babe cesnnssresesecoersernesrnernnnsensee 31 
Injection, water Or COOLANE erro. 66 
Brtbet CATE CEING .ceeeerssossnsssoosrennsesininscrenevetin 121 
Inlet (compressor) temperature vs 
EUG COTSIOO  cissssietincesicrcinsorisenticnionintioton 107 
WOE CUES dictation 16 
Inlet duct, engine air, performance 101 
Inlet Guide VMS oncssesssssssssnseseseennsesseennsee 21 
Inlet SCre@Ms, COMPLeSSOM oncecseessseesnsseeeeeen 18 
Input filter, iGMitiOM .....ccssesssesnsemnssennsseen 61 
INStrUMENTATION, ENGINE ....sssecscccessssesneeeeesee 114 
Insulation blankets .....ssssssssessseeesessesssesnnenn 66 
SP III io sictisernienisnmnncrcemncannntaien 31 
Jet nozzle performance .csscsssssssessseeun 102 
Jee; BHO cee 61 
TPR aie GONG cco tcsvienersntecaencnit 143 
Kelvin temperature ..n.cccosssssssnssssmseseesseeenne 49 
PVN FANE wa aenscscctecssiscteteesrernnone 143 
MRIS 5. ascrckvilasrassinasscmignmemmtncisaateneianie 120 
Laws governing energy exchanges ... 80 
Letters and syMbols o.scccssesssssssnsseeesnseeesen 80 
Limiter, torque (tUrbOProp) neccsesssseeee 140 
ESeaetiE = SMAI PNIE pccsccsccieccssa Secntssa onoecubionenitales 28 
Long reach igniter ...sssssssssssssssssesesnusseesseenean 64 
LOW PreSSUFE COMPFeSSOF -..scsssseessssseeeen 23 
LisIRGINE W..nccinncabndniiepciiepnaencian 145 
Lubrication system, typical ....csccessseseeu 60 
LUbrication SYSTEMS cececsssssssseeesessnssessseeeeeses 59 
Seals rita 3, 79 
Maintenance, gas turbine engine ...... 142 
Manifold, COMpressor aressscsssssseesssssseessesee 20 
PAGES, SN sic seccciccreniesisettaorninien 26 
Manual fuel systern .seseocssssesssssseeeennsssseseeeen 50 
Mass and specific Heat .sreessssssseeeeneen 75 
Mechanical blockage thrust reversers 37 
Mechanically driven accessories ........ 46 
Mice, Exhaust NOZZLE .....sssessessssssssseseneeeee 34 
Military @rmgime ratings .....csssessssseceeeeen 112 
Militcary FUelS asesessscsssccesossceessesesennecssussseesen 143 
Monopropellant fuel o.esessssssssseseeneeeneee 48 





Mounting pads, Crime .neccccececssssssssssseeeen 46 
Multiple combustion chamber ............. 28 
Multistage turbime .....sssssssssssssssssesseseees 31 

Negative torque control, turboprop 70 

Net thrust, defimition occcccsccccsccsssssscsssesseee 8 
Newton’s Second Law wnccccccccscsssssssssssssee 6 
Newton's Thirch LOW o..ceccsssocsssssssssssssssnseesen 5 
Noise levels, Cine ..ercessccsnsssscsssssrsessenee 39 
Noise pattern, CMGiMe o.eceecccssssccsssseesnsssesee 39 
Nose sections, turboprop ...ccscscsssssssssees 15 
Nozzle, adjustable exhaust... 43 
Nozzle, afterburner ...ccccsccsssssssnssssesenseee 43 
Nozzle, COMVErGENE -eccsscccsssscssssssseeeseesesene 34 
Nozzle, convergent-divergent ............... 34 
Nozzle guide VOmes occccssssssssssssssssssssssesseee 31 
Nozzle performance, jet .ccccuscsessen 102 
PI: PII vivricaaee eS a 31 
EN ae ae ae 26 
SIT IIE A icitiaccciarsccnscisiitendiiaiiacensiag 135 
ROE Rg ra ce ee a re 59 
Oil, engine lubricating ....ccccccsssscessseneee: 145 
Oil inlet temperature Gage ...cccccceaecun 116 
Oil pressure GAGE ...ccccsscesenrsseesessnsssneesees 115 
Re I cikeatidtiestuheradaresisainceidebones 59 
Open cycle, CMgime oecccccsccccsssssssnsscsssneseee 78 
Operating cycles, CNGiMe .naccecssessssseeseseen 78 
Operating variable, engine ...... 112, 131 
Operation, turboprop engine ............... 140 
Outer combustion chamber shrouds 29 
Overtempercturiteg ncceoesssscsosssessssesssensesseen 132 
Pads, ENgiN€ MOUNTING ..eccccccsccssnssnnen 46 
Performance curves, turbojet sc 89 
Performance curves, turboprop ... 97, 98 
Performance details, component ......... 101 
Pipe, exhaust (tailpipe) ....ccccscscccssssoeemenn 33 
Pitch stops, turboprop propeller ......... 142 
Pitch, turboprop propeller ......... 139, 140 
Plenum chamber, compressor. occ. 20 
PmeUMatic SHCTtOPs nncessccssccsesrsesrssssseesenesee 47 
Pop-open afterburner nozzle ecco. 44 
Power extraction correction factors... 91 
Power lever, Definition ....ccccccccccccsscccsssssssssee 53 
Power setting curves, turboprop ......... 100 
Power setting procedure, turboprop 140 
Preflight imspection cecssssssssssssssssensesnee 116 
Pressure ratio, AeFiMitiOn .....ccccocsssssesnn 75 

Pressure ratio, engine, use of .. 112, 114 
Pressurizing and dump valve .............. 50 
Pressurizing Valve, Oil ..cccccccssssssossssssessesnnee 60 
Principle of Continuity, the o.oo 81 
Propeller governing range, turbo- 
I sisiiceninnscec catia eS 58 
Propeller governing, turboprop ........... 138 
Propeller governor, turboprop ... 58, 138 
Propeller, turboprop ccccccccssssssccessesesenee 139 
NE I tiniest 50 
Pump, scavenge (Oil) o.ccscecscssssssssssesesen 59 
PID iscascchabasisideablchicnnstinatedbc taal 76 
Ram pressure ratio vs true airspeed 106 
Ram recovery, definition ...cccccccccscccssse- 16 
Ram temperature and pressure rise... 76 
Rankine temperature scale 0.0.0... 74 
ORLA 123 
Ratings, commercial engine ................. 111 
Ratings,engine (general) oss 110 
Ratings, military CNGiMe occcccccscccccseccseeen 112 
Reaction-impulse turbine ....ccccccsceessun 31 
RECCHOM BUTE .....anicsccsncsssssssssnrscccscecevezsees 31 
Reverse thrust, turboprop ....ccccccssossn 14] 
Reverse thrust, Use Of ooocccccccccccssssssssssesee 121 
CS eee 36 
Reynolds number effects ...ccccccscscesen 78 
ROtOr, COMPFeSSOF oaccecccsssssssssessssssssscesseeesees 21 
I 30 
Rpm, disadvantages of using ........... 113 
Rpm for checking thrust ...cccccccsccsen 131 
SCAVENGE PUMPS, OF] oaceecsscocssscsnsesesessseeenes 59 
Screech liners, afterburner o.oo 44 
Screens, compressor inlet ...rccccccccccsuon 18 
Selenium rectifier, ignition occ. 62 
Self-sustaining speed, engine ccc... 47 
Shaft horsepower correction factors, 
I ie schicbinid a dbleemitietanceckcinnl 99 
Shrouded turbine blades 0.0... 32 
Shrouds, combustion chamber .............. 29 
SHUtd Own, CNGIM] ooeccccccccccccsssscsceessessenssseeeee 121 
Silencers, @xHCUSt o.cccccccccccccssscecssssssseeseseeeee 38 
Single axial compressor engine ........ 12 
Single entrance inlet GUct occ 16 
Single-entry COMPFessOT ...ccccccssssssesn 20 
Single-stage tur ime occecscecssssnssssnsessneeee 3] 
SR NII gas socreciocenie scinohisan. 61, 63 
Specific heat amd measss .nreecssosssscssnessnneen 75 
Specific heat, definition ccc 74, 75 


) ) 


= = = = = ES BE RT eSB SS Se ee 

Specification curve, turboprop. ......... 99 
Specification thrust orcs 123 
Speed bias, COMPLESSON -rrceeenrsrnensenenn 24 
Speed bias curve, USC OF cress 131 
Speed bias, explanation of ..... 124, 125 
Speed bias, USC OF -ecrecesnsssensenensenneneen 127 
Speed, SMGIME ececsessemesseeneneeneneenensenenetne 114 
Split COMpPreSsOr ...essscssnseeseessersneenseee 22, 23 
Split tur ime ..sesssssseeessesseeseeneerneensonneeetenen 31 
Spray bars, afterburner -...rcsssscssesseeneen 42 
Stall, COMPressOr ccssssssssssssseesmnnneeeeen 25, 134 
Stall control, COMpPLeSssOr nccseesssreseeseeeen 55 
Standard altitude tables ....ccccssssssseen 104 
Standard day, Adefimition ccc 76 
Starters, COMDUSTION ..ocsssssssssessssssssseeeeeeees 46 
Starters, CMIME nnnnaaseesssscccsssesseeeesssnseseseesssnnnen 46 
Starters, PNEUMATIC ....scsssessesnseeesenseeesen 46 
Starting, SMGiME onnceeessssscssssesssceeesenneeseseennnnesen 116 
Starting SEQUENCE nnacecesssssrsscssesesennsssneesesensers 47 
Static temperature and pressure ...... 76 
Station designation subscripts ............ 77 
Station designations, engine (tur- 

Lape) occcassssssnnoensccconeonecorsnerinevnesoennensnc 10, 84 
Station designations, turboprop  ......... 92 
Stator VANES, COMPFESSOL .....cssssseesenn 21 
Struts, @rctrcast CUCE ocaacsssncsssconccscorsnsossssne 33 
Subscripts, station designation .......... 77 
SUrGe, COMPLESSOM nacceressscsescseesssessenececeenssee 134 
Symbols and letters ..recsccsssseeseeeeeennsenn 80 
Tabs, exhaust nozzle adjustment ...... 34 
Tenth Cerne, OTIS ances ssscosesonccosernvecessonseses 33 
Tailpipe, Cie cacseossssssseensesnnseeennesennnseennse 33 
Tailpipe temperature oncccecccscsecssseseemnnneeen 115 
be: , Sy Sonata eae Neen Ree OE ear 117 
Take-off power check curves, turbo- 

OP sc sessiehosisieennasinarstnantae cinemas 140 
Take-off thrust check curve, tur- 

gy Neecoeie eat Se oielees ieee 126, 127 
Take-off thrust setting curve, turbo- 

fe RRR a a ee 127 
Temperature and heat ..cccssssccssseessseessenn 73 
Temperature conversion table... 105 

Temperature effect, dual compressor 24 
Temperature effect on horsepower, 
net thrust and fuel flow (turbo- 
PI ON Yassin sassiecninnatrren 97 


Temperature Limits ....rccceseesssnsesneesesneessen 132 
Temperature ratio, definition ............. 76 
curves, use Of ..... 125, 126, 127, 131 
Temperature SCOIES ...ccccssssasemseeneenernsenn 74 
Temperature variation with altitude 3 
TPIAR IRR asiscvneeceecirtinniccccenectmncrcteercins 7 
Thermodynamics, Dcisic ccsesssoessesenseensen 80 
Thermodynamics, First Law of ........... 80 
Thermodynamics, Second Law of ...... 80 
Thermometer SCOI@S esccsssssssssssseeseeseeseemnsssee 74 
FON ei ceetancreinnpenieannacnnament 77 
Tharcotthes, LeFTTRIIA osecncaccscssenessosionenssiovtensenen 53 
Thrust COCFCIOIE -nnccsssccsssnnsscscsecssorecossernnnnsne 90 
Thrust correction factor for duct loss 92 
Tharcih, COTATI icecrccicicesnecensssicsigioni 7 
Thrust, development OF -ccrcsssssssemssseeee 5 
THAPUSE, CEOS icsccereacssssnsonosooernnencercnnnnoroennness 9 
TRIN, GIN srciiccicsesccrreeiceemeseintemanncsio 8 
TEAFRIGE POV AEGIS accssccscssissteinieesscsssarecsrnesinnenest 36 
Thrust specific fuel consumption ......... 10 
Torque limiter, turboprop oases 140 
Torquemeter constant, turboprop ...... 14] 
Torquemeter, turboprop. ........... 114, 140 
Total temperature and pressure ... 76, 77 
Trigger Spark, iQritiOr ...ccsssssssssmssseeenen 63 
TriM CUFVE, ENGINE woieecccsssssessseecresees 126, 127 
Trim speed, Definition ...eercsssssssseenneeenen 123 
Trim speed, field Use OF .rcccsssssseneeen 131 
Trimming dual axial compressor en- 
RIVE aeccensessvsereinsnnesssossnsensonrvrannnneosvvcsennsseoseenion 123 
T rit, CP ie™ cessssssssssescsceeesseeneeeeeens 34, 123 
Trimming, part throttle method ............ 129 

True airspeed vs compressor inlet 
TOMPCLAtULe nccscsersssessseeseerseceesnnsssenensnossone 107 

True airspeed vs ram pressure ratio 106 

Turbine discharge pressure measure- 

NE is a cceatancagenonnnine 33 
Turbine discharge pressure, use 

SRO Re eS 112, 114 
Turbine discharge temperature 115, 132 
TUrRSe RE wii 3] 
Turbine performance eeecsscssssssssseeeaeeeueeen 102 
TREE ccckcctceieeninaine 30 
Turbofan @mgies ecssssssssssssssessensecssnsse 13, 14 
Turbojet engine performance analy- 

cass imsasorpostbepthbibeocciniannininaisieinsapistaniesin 84 






Turboprop blade 
ANTES nino 137, 138, 139, 140 
Turboprop characteristics ...cccesssscssenseee 136 
Turboprop compared with turbojet ... 137 
Turboprop compressor inlet .....cccscoo 18 
TUrbOprop CM gie eeeescsssssesscsesessneseseceesenneee 13 
Turboprop @Ngine CUFVES ...ccscccscseneenn 98 

Turboprop engine performance anal- 
SOTTO scisccansichesiebin ceili 92 
Turboprop fuel and propeller control 58 

TUrboprop Operation ..eecscccsssccsssssseeseen 140 
Turboprop performance... 137, 138 
Turboprop power setting curves ........ 100 
Turboprop propeller ..ccccccssssssssssssnesen 139 


Turboprop specification CUFVE .....ccc... 99 
Two-position exhaust Nozzle .....ccsc... 43 
Types of gas turbine engines .............. 11 
ee Ce eR 30, 31 
Variable-area exhaust nozzle ........... 34 
Waiter ijectionn .ncccccccccsssssvoossnsscssnnesssceseee 52, 66 
Water (injection) purity ....ceccccscossssoeen 68 
Water injection system .....cccccsasssssssees 68 
~ Weight, definition .......ccccccssssssssnsesussesnee 75 
Wheel, turbine iccccssssssssccsssscscesseeee ~ oe ae 
Wirrcmillimgg Arc sesssssesssscssssnesesnsseenneee 70 


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