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for the

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:1 PRE f h!WXARY RESULTS OF AN ALTITUDE -WllW’TUNNEL Lst

9

4;’ INVESTIGATION OF A TG-100A GAS 3 ‘Q .

1: ts$ Q 7 i

TURBI?iE-PROPELLER ENGINE ‘L * :*

\$ ;p .- COMBUSTION-CHAMBER CHAkACTERISTICS

3: _ --

r Robert M, Gexsenheyner, and Joseph J. Berdysz

Flight Propulsion Research Laboratory Cleveland, Ohio

CLASSIFIED DOCUXENT

This document contains classified information affec- nited States within ,UX 50:31 and 32. of its contents in

is prohibited by be imparted only

Services of the officers and em-

ent who have a legit- imate interest therein, and to United States citizens of known Lo be informe l

alty and discretion who of necessity Dust thereof.

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

WASHINGTON ‘. )’ I. 1 . ,; _1 s--. . .._- r, * _

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cl “i”~~ll~~~~~~~~~~lll~ lj&CA RM No. SE7LO9 -----

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-

NATIONAL ADVISORY COMMlTTEE FOR AWONAUTICS

RiZXARCHMEMORAIDUM

for the

Air Materiel Command, Army Air Forces

l?RELIMINARYRESULTS OFANALTITUDE-WIND-T

INVESTSGATSON OF A TG-1OOA GAS

TURBINE-PROPELLER ENGlX$

Y- COMBUSTION-Cm CHARACTERISTICS

By Robert M. Geisenheyner, and Joseph J. Berdysz

An investigation to determine the performance and operational characteristics of the TG-1OOA gas turbine-propeller engine was conducted in the Cleveland altitude wind tunnel. As part of this investigation, the combustion-chamber performance was determined at pressure altitudes from 5000 to 35,000 feet, compressor-inlet rm-pressure ratios of 1.00 and 1.09, and engine speeds from 8000 to 13,000 rpm. Combustion-chamber performance is presented as a function of corrected engine speed and.correcte& horsepower.

For the range of corrected engine speeds investigated, over- all total-pressure-loss ratio, cycle efficiency, ana the frac%ional loss in cycle efficiency resulting fram pressure losses in the combustion chambers were unaffected by a change in sltitude or compressor-inlet ram-pressure ratio. The scatter of combustion- efficiency data tended to obscure any effect of altitude or rsm- pressure ratio. For the range of corrected horsepowers investi- gated, the tot&L-presstie-loss ratio an& the fractional loss in cycle efficiency resulting from pressure losses in the combustion chambers decreaaecl with an increase in corrected horsepotjer at a constant corrected engine speed. The combustion efficiency remained constant for the range of corrected horsepowers investi- gated at all corrected engine speeds. 0

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1 ._. - . I -. . . I (1

2 NACA RM No. SE7LO9

INTRODUCTION

An investigation to determine the performance and operational ch&racteristics of the TGlOOA gas turbine-propeller engine was conducted in the NACA Cleveland. altitude wind tunnel at the request of the Air Materiel Comma&, Army Air Forces. The performance characteristics of the component pads of the engine were determined. in addition to the evaluation of the over-all performance chasacter- istics. Various phases of the investigation sxe reported in refer- ences 1 to 4.

The combustion-chamber performance is presented herein to show the effect of engine operating conditions on combustion-chamber over-all total-pressure loss, combustion efficiency, cycle effi- ciency, and the fractional loss in cycle 'efficiency due to combustion-chamber pressure losses. Data were obtained for pres- sure altitudes from 5000 to 35,OOu feet, compressor-inlet ram- pressure ratios of 1.00 and 1.09, and engine speeds from 8000 to 13,000 rpm.

WZX!RIFTION OF COMBUSTION CHAMBEX? '

The TGlOOA gas turbine-propeller engine has nine cylindrical counterflow combustion chambers (fig. 1). The air leaving the last stage of the compressor is turned 180° before entering the combus- tion chambers and passing into annular spaces between the casings an& the liners. The casing (fig. 2(a)) is 6 inches in diameter and 14$ inches long and contains a removable liner (fig. 2(b)) that divides the combustion zone from the passsge for the entering air flow. A series of- openings in the wall of the liner allows air to pass into t&e combustion zone, where it is mixed with fuel sprsyeh from an atomizing nozzle located in the center of the combustion-chamber dome. The fuel is ignited by spark plugs located in the dome of two of the combustion chambers; fuel in the other combustion chambers is ignited through cross-fire tubes.

ENGINE INSTALLATION AND INSTRUMEXLWION

'The TGlOOA engine was installed in a streamlined wing nacelle, which was mounted in the 20-foot-dismeser test section of the Cleveland altitude wind tunnel (fig. 3). A sections3 drawing of the engine showing the locatiop of measuring stations is given Pn ffgure 4. A complete description of the engine, the'installation and the instrumentation is presented in reference 1.

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NACARM No. SE7L09 3

The instrumentation used in this analysis was located at the compressor inlet (station: 2), the compresBor outlet (station 3), the turbine inLet (station 5), and. the tail-pipe-nozzlG outlet (station 8). (station 4),

Because the instrumentatiqn at the compressor elbow, was inadequate for determining pressure losses across

the combustion chambers, the instrumentation at the compressor out- let (station 3), was used.

The compressor-outlet instrumentation consisted of three rakes, located 120° apart, with three total-pressure tubes and two thermo- couples each. The location of all instrumentation at this station, including five wall statiotubes and two probe static tubes, is showndn figure 5. Turbine-inlet instrumentation consisted of five total-pressure tubes and five wall static tubes located as shown in figure 6. The tail-pipe rake at station 8 (fig. 7) contained 16 total-pressure tubes, three static-pressure tubes, and six thermo- couples ana was located with the plane of measurement 1 inch upstream of the tail-pipe-nozzle outles.

PROCEUURE

The investigation was conducted at pressure altitudes ranging from 5000 to 35,000 feet and at compressor-inlet ram-pressure ratios 0f 1.00 and 1.09. At each altitude and compressor-inlet ram-pressure ratio, the engine speed was varied from 8000 to 13,000 rpm. The engine power measured at the torquemeter ranged from 70 to 1050 horse- power. Ambient temperatures were maintained at approximately NACA dea altitude conditions. Average vaLuea of pressure and tem- perature used in this analysis were taken from references 1 and 3.

SYMEQIS

The following symbols me used in this report:

A cross-sectional area, square feet

cP specific heat of gas at constant pressure, Btu per pound OR m

cP,J average specific heat at constant pressure from station 9 to station 0, Btu per pound OR

f fuel-air ratio

I I___ __~ _._ .

v .-.- ~----_ _._ _ _.

I - ~:‘~,. _

4 l'IACA RM No. SE7LO9

Q acceleration due to gravity, 32.2 feet per second per second

ghP horsepower loss in high-speed reduction gesr

H enthalpy, Btu per pound

Ha enthslpy of air, Btu per pound

Hf enthalpy of fuel, Btu per pound

hf heating value of fuel, 18,700 Btu per pound

- .

hp shP +ghp

J mechanical equivalent of heat, 778 foot-pounds per Btu

N engine speed, rpm

P totsl press&e, pounds per squsxe foot absolute

@d/p3 over-all total-pressure-loss ratio across ccmbuation chambers due to friction and plomentum changes

, pounds per squsre foot

P

R

shp

T

Ti

t

wa

Wf

wt3 Y

static pressure, pounds per square foot absolute

gas constant, 53.4 foot-pounds per pound OR shaft horsepower measured at torquemeter

total temperature, OR

indicatea temperature, OR

static temperature, 41

air flow, pounds per second

fuel flow, pounds per hour

gas flow, powa per second

ratio of specific heats for gases

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., _ .’ .‘, i’--~

~~.. _

~ _.-.p G,‘. ,C “.

I’. <. 8 -’

NACA RM No. SEX09 5

s ratio of compressor-inlet total pressure to static pres- sure of NACA standard atmosphere.at sea level

8 ratio of compressor-inlet absolute total temperature to absolute static temperature of NACA standard atmosphere at sea level

tt cycle efficiency

cycle-efficiepcy loss

fib combustion efficiency

Subscripts:

0 tknel-test-section free air stream

2 compressor inlet

3 compressor outlet

5 turbine inlet

8 tail-pipe-nozzle outlet

9 station in jet where' static pressure*first reaches free- stream static pressure

The data were generalized. to NACA standsrd sea-level conditions bi the following parameters:

WqF corrected engine speed, rpm

hp/@@) corrected horsepower

LN~ZI'HODS OF CALCULATION

Gas flow =a air-flow. - The gas flow at the tail-pipe-nozzle outlet was obtained from

,

6 NACA RM No. SE7LO9 '

The air flow was then determined from

wa = wg - (Wf/3600)

Turbine-inlet temperature. - Turbine-inlet temperature was calculated framthe enthslpy drop through the turbine and the enthalpy at the tail-pipe-nozzle outlet:

H5 =(v]+ k3 -Hz)tH8

Hs T5 =- cP,5

An integrated value of cp 5 was use‘& in this equation. The value of ghp used in ca3culatikg H5 was estimated to vary from 50 horsepower at an engine speed of 13,000 rpm to 25 horsepower at 8000 rpm.

Jet-gas temperature. - The gas temperature in the jet was determined from temperature and pressure measurements at the tail- pipe-nozzle outlet. For a thermocouple recoveryYfactor of 0.85, the total temperature at the nozzle outlet was calculated from .

733-l

1 + 0085

If the.total pressure, the total temperature, and the ratio of specific heats at the nozzle outlet (station 9),

(station S), an& in the jet, are assumed to be equal, I

Y&-l

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- ‘. -. : ,,-

“T, ----Yy--: _, r ..‘ ..‘,.,.+‘. :,

:

NACA RM No. SE7LO9 7

MEKCHODS OF ANALYSIS

Combustion efficiency. - Combustion efficiency is defined as . the ratio of the actua$ increase in ent&lpy of the gas to the theoretical increase that would result from ccmplete combustion of the fuel charge:

Tb =, tHa,SmHa,3 ) + (rr, s-Q,+

rkf

Cycle efficiency. - Cycle efficiency was detemnined according to the standard. definition

9 = heat supplied by source - heat rejected. to sink heat supplied by source-

tl = (Hs-Hd - Cp j(tg-to)

H5-'H3

Loss in cycle efficiency. - The expression for loss in cycle efficiency resulting from pressure losses in the combui3tion chamber was calculate& by the method given in reference 5:

AT) = HS'H3

where y is the average value between stations 5 and 9.

RESULTS AND DISCUSSION

The presentation of combustion-chamber performance includes the evaluation of over-all loss.in total pressure across the combustion chambers, ocmbustion efficiency, cycle efficiency, and loss in cycle efficiency resulting froqpressure losses in the combustion chambers.

The over-all total-pressure. loss across the combustion chambers may be considered as the sum of two comppnents: the loss due to friction and that due to the addition of heat by combustion, or

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_ ’ . I;

8 NACA RM No. SE7L09

momentum-pressure loss. Reference 5 gives a method of determining these components by assuming that all the friction loss occurs before the gas is heated. This assumption is valid fop a through- flow type of combustion chamber. In a counterflow combustion chamber; however, the assumption is invalid because preheating occurs during the flow of the entering air along the outside of the liner or basket. Consequently, the method of reference 5 is not applicable to the counterflow type of combustion chamber ana only over-all pressure losses are considered in this report.

Pressure-Lobs Rqtio

Over-all total-pressure-loss ratio (MT)/23 for the range of corrected engine speeds investigated is presented in figure 8(a) for altitudes from 5000 to 35,000 feet at static conditions and a tail-pipe temperature of approximately 1500° R. A similar compari- son of pressure-loss ratio is given in figure 8(b) for compressor- inlet ram-pressure ratios-of 1.00 and 1.09 at-an altitude of 25,000 feet. Changes in altitude or ram-pressure ratio did not appreciably affect the total-pressure-loss ratio across the com- bustion chamber. The total-pressure-loss ratio was 0.025 for the range of corrected engine speeds investigated. The variation of total-pressure-loss ratio with corrected horsepower is shown in figure 9 for various corrected engine speeds at an altitude of 5000 feet ma static conditions. At a constant corrected engine speed, the total-pressure-loss ratio decreased with an increase in corrected horsepower. For a constsnt corrected horsepower, the total-pressure-loss ratio increased as the corrected. engine speed- increased..

' Combustion Efficiency -

The relation of combustion efficiency to corrected engine speed is presented in figure 10(a) for altitudes from 5000 to 35,000 feet at a ram-pressure ratio of 1.00 and in figure 10(b) for ram-pressure ratios of 1.00 and 1.09 at an altitude of 25,000 feet. The scatter of these data tended to obscure any effect of altitude or ram-pressure ratio on combustion efficiency. The variation of combustion efficiency with corrected horsepower is shawn in figure ll for vsrious corrected engine speeds at an altitude of 5000 feet-and static conditions. For the range of corrected horsepowers investigated., varying the corrected engine speed had no appreciable effect on combustion efficiency. Values

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NACA RM No. SETLO 9

of combustion efficiency over 1.00 sre considered the result of inaccuracies in measurement of turbine-outlet temperature and shaft horsepower.

Cycle Efficiency

Cycle efficiency as a function of corrected. engine speed at a tail-pipe temperature of approximately 1500° R is shown in figure 12(a) for altitudes from 5000 to 35,000 feet at static conditions and in figure 12(b) for ram-pressure ratios of 1.00 and 1.09 at an altitude of 25,000 feet. As the corrected engine speed was increased from 8170 to 14,18oLrpm, the cycle efficiency was raised from approxi- mately 0.04 to 0.17. Changes in sltitude or ram-pressure ratio had no apparent effect on cycle efficiency. The variation of cycle efficiency with corrected horsepower at various corrected engine speeds and static conditions is shown in figure 13. Cycle effi- ciency increased with increasing corrected horsepower from 0.03 to 0.16 over the range of the investigation. At a constant cor- rected horseptier, a change in engine speed had no effect on cycle efficiency.

Loss in Cycle Efficiency

Cycle-efficiency losses that resulted from combustion-chamber pressure losses are presented as fractional loss in cycle effi- ciency (Aq)/?j for the range of corrected engine speeds investi- gated at sltitudes from 5000 to 35,000 feet and static conaitiom at a tail-pipe temperature of approximately 15OOO R (fig. 14(a)). Fractional loss in cycle efficiency decreased with an increase in corrected engine speed over the entire range'of the investigation. The change in altitude had no appreciable effect on (Aq)/q for the range of corrected engine speeds. Increasing the ram-pressure ratio fram 1.00 to 1.09 reduced the fractionad, loss in cycle effi- ciency approximately 0.01 for the range of corrected engine speeds investigated (fig. 14(b)). The fractional loss in cycle efficiency for a range of corrected horsepowers at an altitude of 5000 feet and static conditions is shown in figure 15. For a constant cor- rected engine speed, (67)/q decreased with sn increase in cor- rected horsepower. At a constant corrected horsepower, (Aq)/? increased with an increase in engine speed. ,

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.-

10

.

IWCA RM No. SE7LO9

SWWY OF RI&STILTS

An investigation of counterflcw combustion chambers operating in a TGlOOA gas turbine-propeller engine over a range of pressure altitudes from 5000 to 35,000 feet and. ram-pressure ratios of 1.00 an& 1.09 gave the following results:

1. Total-pressure-loss ratio was unaffected by changes in altitude at a constant tail-pipe temperature and remained at a value of approximately 0.025 for the range of operating conditions investigated. A change in ram-pressure ratio from 1.00 to 1.09 had no appreciable effect on total-pressure-loss ratio. At a constant corrected engine speed, the total-pressure-loss ratio was reduced as the corrected horsepower increased.

2. The scatter of data tended to obscure any effect of alti- tude or ram-pressure ratio on combustion efficiency. Varying the corrected horsepower at a constant corrected engine speed had. no effect on combustion efficiency.

3. At a constant tail-pipe temperature, the cycle efficiency increased with increasing corrected engine speed, but was unaf- fected by a change in altitude or ram-pressure ratio. At a constant corrected. engine speed, the cycle efficiency increased. as the corrected horsepower increased. At a constant corrected horsepower, however, char&es in corrected engine speed had no effect cn cycle efficiency.

__-..-_ -_ . _, ,: ;,‘.~, ; (__. .- ; .’ ‘. :- , . . ,y - :.:, :,,*r ._

:

NACA RK No. SE7L09 ll

. 4, Fractional loss in cycle-efficiency, the result cf pressure

losses in the combustion chambers, decreased with an increase in corrected engine speed and was not appreciably affected by a change in altitude or ram-pressure ratio. At a constant corrected engine speed, the fractional loss in cycle efficiency decreased with increasing corrected horsepower.

Flight Propulsion Research Laboratory, National Advisory Committee for Aeronautics, ,

Cleveland., Ohio, November 25, 1947.

Robert M. Geisenheyner, Mechanical Engineer. .

Approved:

1.

2.

Alfred W. Young, Mechanic&L Engineer.

Abe Silverstein, Aeronautical Engineer.

REZ%REJKFS

Saari, Martin J., end Wallner, Lewis E.: Preliminary Results of an Altitude-Wind-Tunnel Investigation of a TG-1OOA Gas Turbine-Propeller Engine. I - Performance Characteristics. NACA RM No, *E7FlOa, Army Air Forces, 1947,

Conrad, E. W., and Durham, J. I).: Preliminary Results of an Altitude-Wind-Tunnel Investigation of a TG-1OOA Gas Turbine- Propeller Engine. II - 1&ximilling Characteristics. NACA RM No. E7G25, Army Air Forces, 1947.

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._ I _ ‘.. ,.

12 NACA RM No. SE7LO9

3. Geisenheyner, Robert M., & /Berdysz, ioseph J.: Preliminary Results of an Altitude-Wind-Tunnel Investigation of a T-G-1OOA Gas Turbine-Propeller Engine. III - Pressure and Temperature Distributions. INWARM No. E7J02, Army Air Forces, 1947.

4. Wsllner, Lewis E., and Saari, Martin J.: Preliminary Results of an Altitude-Wind-Tunnel-Investigation of a TGlOOA Gas Turbine-Propeller Engine. IV - Compressor and Turbine Performance Chsxacteristics. NACARM No. SE7J20, Army Air Forces, 1947.

5. Pinkel, I. Irving, and Shames, Harold: Altitude-Wind-Tunnel Tests of the General Electric TG-180 Jet-Propulsion Engine. VI- Combustion-Chamber Performance. NACA RM No. E6Xl7, Army Air Forces, 1946.

C :

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Figure 1. - Side view of TG-IOOA gas turbine-propeller engine showing installation of counterf low combust ion chambers.

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NACA RM No. SE7 LO9 I5

- _ _-. ~~ ..-~.--

(a) Casing with spark plug installed.

C- 19318 .I- -. -- 8- 12-47

(b) Liner.

Figure 2. - Counterflow combustion chamber of TGIOOA gas turbine- propel I er engine.

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-. , . < ._ : : I -’ ^-

I -.

1

_., :- _c _ -:.-----~-- - --- , . r: ; -..; ‘- :: ~ .; -: _’ _ - -.:.. I -. _’ ..I _. ’ *. <_ . *-. -_ .

NACA RM No. SE7LO9 17

Figure 3. -- Front view of TG-IOOA gas turbine-propeller engine in alti- tude wind tunnel.

-.-~ ..- --_~ -.-- _----- ..-._.- -- --., ---- .,-..- - __.--_.*_y_ ___ - . -.._-----.*.-*- ------ _.-- _ -.--

. ~.

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> -. ; c I . . & , -. ,

._ ~.-- ;

-_ : I .- . .

Station

2 Compressor inlet 3 4

Compressor outlet Corn ressor elbow

5 Tur ine inlet Ii 6 Turbine outlet. 7 Exhaust-cone outlet 8 Tail-pipe-nozzle outlet

I /Propeller

Station 2 7

Exhaust cone

I 8

Pe

Figure 4.. - Side view of TG-IOOA engine showing location of measuring stations.

NACA RM No. SOL09

Y -

OTotal-pressure Wbe l Static-pressure tube c9 Thermocouple

Figure 5. - Location of instrumentation at compressor outlet, lodking aft, station 3.

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NACA Rh4 No. SE7LO9 21

o Total-pressure tube 0 Static-pressure tube

so l

50 l

00

4 <

s50 ’

00

$l$$ y A ’ \ \

\ \\

\i+) -- 0 - - -I

.

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YifI4-9 . Figure 6. - Locat ing aft, ion of instrumentation at turbine inlet, look

station 5.

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NACA RM No. SEi’L03 23

b

--. _ . - ._ r .- __-. -----------l~~f~-~_~-.,.~~f_ ---.__- --.-. __--c-^-_-_- --__ ~_ _ ___. - - . . . . . : 1,

_. . .

Figure 7. - Instrumentation at tail-pipe-nozzle outlet, station 8.

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Altitude ml I I I I I I I I I I

0 ‘5,000 .

q .04 15,000 0 25,000 A 35,000

0 " I7 I 0 m n r I3 .02 0 U 8 A .A

(a) Altftude varied; compressor-Inlet ram-pressure Patfo, l&00

.04

602

=T.pi&Y

' Ov 8 9 10 11 1% 13 b4 16 x IO"4 Corxw&ed engine speed, HA*

(139 compressor-inlet ran-prosswe m tfo vaxpbed; albtitub, 25&HJO iZW&

Figure 8. - Effect of corrected engine.speed on over-al I total-pressure-loss ratio for various alti- tudes and compressor-inlet ram-pressure ratios: Tail-pipe temperature, 1500’ R.

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f . is 2 ? 1 r(s 2 ’; % ;J’ F j i I r- l $ 4 8

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Correc ted

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1 0 8

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O O 2 0 0 4 0 0 8 0 0 8 0 0 1 0 0 0 1 2 0 0 1 4 0 0 Uorrec ted ho rsepower , h p /(@ @ )

, - E ffect of co r rec ted h o r s e p o w e r o n over -a l I to ta l -pressure- loss rat io for va r ious cor rec te e n g i n e speeds . A It i tude, 5 0 0 0 feet; c a m p resso r- i n I et r a m - p ressu re rat i o, I .O O .

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(a) P%%%hade vaz?%ed; oomprsssor-inlet ham-pmssum mtfo, LOO, 4 I I t 6 c % Lad@

g 2 s P Q locm v

080 \ ‘p 8 9

corr::eea .i?@gL3 speod,‘$w

b9 Compm6aorQnlet IP~-~PBGG~~ rat80 vmBe Figure IO. - Effect of corrected engine speed on combustion

camp resso r- I n I et ram-p ressu re rat i 0s. (Tai led points ind

,d; ahL%tude, 25,OOO foe&o efficiency for various altitudes an

icate tai I-oi9e temgerature of 1500’ d

R. 1 Y

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l&o b 0 on n

g 0 4 t:

KIijpJ7

200 400 I

600 8OQ 1000 Uomee ted hp/(tj*)

1800 1400 horsepower D

Figure II. - Effect of corrected horsepower on combustion efficiency For various corrected engine cl

speeds. Altitude, 5000 feet; compressor-inlet ram-pressure ratio, I .OO. ,-

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z

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P . z 0 e

m

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(a) Altitude varied; compressor-inlet ram-pressure ratio, 1.00. 1 I I I I I I I I I I I I I

Uomsoted en&x speed, W/s ’ (bsg cmp~essoxJ-%wls& Fam-gress~ mtio war%ed; al&8tude, 25,000 fee%0

Figure 12. - Effect of corrected engine speed on cycle ef’ficiency for various altitudes and camp resso r-i n I et ram-p ressu re rat i 0s o Tail-pipe temperature, 1500° R.

, .

,

Corrected engine

swybJ*

0 13,286

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I I

.

'0 0 200 400 600 800 1000 1200 1

Corrected horsepower, hp/(Ss)

s Fe 2 z 0 .

v, Figure 13. - Effect of corrected horsepower on cycle efficiency for various corrected engine speeds. 3

Altitude, 5000 feet; compressor-inlet ram-pressure ratio, I .OO. 5

B T;” 6 8 VI 0 ii “I Q)

E

k

UY (a) Altitude varied; compressor-inlet ram==press\uPe ratio, LOO,

Figure I

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Ram- 020. pressure

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010

. 0

7 a 9 10 Corrected &he spe%, N/q

(b) Compressor-Inlet ram-pressure ratio varied; altitude, 185,000 feet, 4. - Effect of corrected engine speed on loss in cycle-efficiency ratio for various altitudes

and comprsssor-inlet ram-pressure ratios. Tai l-pipe temperature, 1500’ R. w

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Corrected horsepower, hp/(6@) Figure 15. - Effect of corrected horsepower on loss in cycle-efficiency ratio for various corrected

eng i ne speeds. Altitude, 5000 feet; compressor-inlet ram-pressure ratio, 1.00.

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