ACE No. L5EI7
NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
WARTIME REPORT ORIGINALLY ISSUED June 19^5 as Advance Confidential Report L5EI7 PEEFOBIANCE OF C0MPRESS0R-1UEBI1IE JET-PROPULSION SYSTEMS By Carl B. Palmer
1
Langley Memorial Aeronautical Laboratory Langley Field, Ta.
MAC A WASHINGTON NACA WARTIME REPORTS are reprints of papers originally issued to provide rapid distribution of advance research results to an authorized group requiring them for the war effort. They were previously held under a security status but are now unclassified. Some of these reports were not technically edited. All have been reproduced without change in order to expedite general distribution.
L - 27C
L. R - / 7 / £
3 1176 01354 2411
NAOA ACR No. L5E17 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS ADVANCE CONFIDENTIAL REPORT PERFORMANCE OF COMPRESSOR-TURBINE JET-PROPULSION SYSTEMS By Carl B. Palmer SUMMARY
An analysis of the performance of compressor-turbine Jet-propulsion sy&tems was carried out by calculating the thrust power fron a compressor-turbine jet engine with a systematic variation of pressure ratio, fuel-air ratio, compressor and turbine efficiencies, flight speed, altitude, and maximum gas temperature. Increasing the compressor and turbine efficiencies from 70 to GO percent was found to double the over-all efficiency of the engine at 300 miles per hour (iül0 fps)« Increasing the speed from 300 to 6C0 niles per hour (880 fps) increased the over-all efficiency by 7 to 10 percent. The maximum power output at a particular altitude was shown to tn approximately proportional to the bemperature difference between the combustion chamber and the free atmosphere. INTRODUCTION The basic principles of thermal-air Jet propulsion have long been understood but not until recently have systems been devised that are capable of applying these principles to the propulsion of passenger-carrying airplanes. Th« method that appears to have the greatest potentialities makes use of mechanical compression of atmospheric air and continuous burning of fuel in the compressed air. One of the early practicable systems, the Italian Caproni-Campini, made use of an ordinary internal-combustion engine for running the compressor.
KACA ACR No. L5E17
This system eliminated the propeller but still had the heavy weight and the complication of the reciprocating engine combined with the low efficiency of a marginal jet-propulsion system. The use of a gas turbine for driving the compressor offered the advantages of simplicity and low engine weight per horsepower output, but the thermal efficiency was impracticably low because the turbine had to operate at low temperatures to prevent blade damage. Brown, Boveri & Company, Limited, had developed practical gas-turbine power plants for stationary installations, in which weight was no problem and a considerable amount cf regeneration could be used. The thermal-air jet engine with turbinedriven compressor beca? :? a practical means of aircraft propulsion, however, only with the development of materials for pas-turbine rotor blades that could operate continuously at temperatures of 1200° P or higher and the development of a lirht-weight rotary compressor capable of 3iving a pressure ratio of at least 3* at greater than 60 percent efficiency. Reference 1 describes a turbinecompreesor unit suitable for use in a jet engine. Although the temperatures and efficiencies at which the turbine-compressor Jet engine becomes practicable are of interest, it appears more important to inquire into the effect of further improvement in temperatures, efficiencies, and other pertinent factors in Jet-engine performance. An analysis of the effects of various jetengine design and operational parameters may indicate the most profitable lines of developmental research and the amount of Improvement in performance and efficiency to be expected from such research. Such an analysis of jetengine performance is presented herein. The altitude and speed of flight, the turbine and compressor efficiencies, the fuel-air and pressure ratios, and the combustionchamber temperature are varied to show the effect of each not only on the engine thrust power but also on the optimum values of the other parameters. Among the pertinent topics not considered in this analysis are engine weights and the effect of arbitrarily changing the fuel rate for a particular power olant.
NAGA AGR No. L5E17
METHOD'OP ANALYSIS The calculations for this study were made on a Mollier chart for air (see fig. 1, which was transformed from a chart in reference 2 and is placed at the end of the report) by the methods described in reference.5. For purposes of this analysis, air compression by ram is isentropic, mechanical compression is at an arbitrarily assigned efficiency, cmbustion takes place at constant pressure and a given efficiency, energy is taken from the working fluid by the turbine at an arbitrarily assigned efficiency, and the air after passing the turbine accelerates isentropically to free-stream static pressure. For every combination of altitude, speed, and maximum allowable temperature, various pressure ratios are used; in eaoh case the amount of fuel required to raise the air temperature to the defined maximum is burned. For each set of conditions two combinations of fuel-air ratio and pressure ratio are stressed - one giving maximum power and one giving EiaxiTrom over-all officiency. The calculations on the Mollier chart indicate the thrust power from each pound per second of conducted air. Tn OT«der to show more cüearly the effects of altitude and speed, the design is asrumed to be euch that the weight flow of charge air Is proportional to the free-stream stagnation density; the base used Is ).j.O pounds per second at öOO feet per second at sea level. (See fig. 2.) This assumption is in reasonable accord with results from actual installations because the air velocity should be approximately constant In the engine in order to maintain the compressor and turbine efficiencies. Whenever possible, the graphs of results are drawn with scales of both power and power per (pound per second) of charge air. The graphs for zero flight speed at sea level show the static thrust force; for all other conditions, power rather than thrust is'shown. The compressor and turbine are operating at defined efficiencies so that, when the fuel-air ratio and the compression ratio are changed at a particular altitude and speed, the curves represent an Infinite number of engines, each of which is designed to have the defined efficiencies at the particular operating conditions under consideration.
NACA ACR No. L5E17
hQ
Altitude (ft)
ho Sea level
=52 o 2). ^•^^"^
25,000 1
16 -—q.0,000 _
1
i
—50,000 -60,000 .
200
l;.00
600
800
1000
V0, fpa Figure 2.- Woight flow of charge air. The results of the analysis are oresented In two parts. In the first part only the jet engine is considered, without reference to any airplane In which it mi.cht he Installed, and in the second part the performance of a particular Installation is discussed. The symbols used herein are defined in appendix A, and the conditions and assumptions used are discussed In appendix B.
NACA ACR No. L5S17
PERFORMANCE OF" JET ENGINE Figure 3 presents a set of cycles on the Molller chart for the purpose of Illustrating the effect of pressure ratio and fuel-air ratio on the thrust from 1 pourid of air. The vertical distances (enthalpy changes) In cycle B are- significant In the following manner: The distance 0 to 1 Indicates the velocity with which the air approaches the engine, 1 to 2 shows the energy added by the compressor, and 2 tc 3 shows the energy added by burning fuel at constant pressure. At a particular altitude, flight speed, compressor efficiency, and maximum temperature (tempcratuve at point 3)» the location of point 2 uniquely determines the fuel-air ratio and the pressure ratio, so that one ratio may be plotted as a function of the other. The distance 3 to I4. shows energy taken out by the turbine, and ij. to 5 indicates the exit velocity of the propelling Jet. The distance 5 to 5' is the same as 0 to 1 so that, when point ij. falls on 5», the
o •p
cd
u o &
o
•P
OS ft
Locus of Ij. Locus of 5»
Ö
Entropy Figure-3«- Jet-engine» cycles.
NACA ACR No. L5E17
exit velocity equals the approach velocity and the conducted air contributes neither thrust nor drag. The thrust Is, therefore, determined by the distance J4. to 51» which In conjunction with 0 to 1 shows the velocity Increase of the conducted air; that Is, Thrust cc AV cc distance I4. to 5 - -\Ai stance 51 to 5 Cycle A, which has hiph pressure ratio and low fuel-air ratio, and cycle D, which has low. prespure ratio and high ftiel-air ratio, show little or no thrust. Cycles B and C give about equal thrust; and maximum thrust would be obtained with a cycle intermediate to B and C. Figures 1). to 6 show the variation of thrust power with fuel-air ratio for various turbine and compressor efficiencies. The pressure-ratio curve Is also shown as a function of fuel-air ratio. In order to find the pressure ratio for a particular point on a thrust-power curve, read the value of R for the fuel-air ratio corresponding to the point on the thrust curve. With the maximum temperature fixed, operation is possible only in a narrow range of fuel-air ratio and pressure ratio. These fuelair and pressure ratios are shown for three gas temperatures In the following table: Maximum temperature '"ressure ratio
(°P)
Puel-air ratio
1500
5 to 6
1800
8 to 9
.019 to
.016
2100
10 to 12
.022 to
.020
0.015 to 0.013
In this table both turbine and compressor efficiencies are about 80 percent. If these efficiencies were 70 percent, the fuel-air ratios would be 0.001 to 0.002 higher, and the pressure ratios would be about two-thirds of those shown. Decreasing the turbine and compressor
NACA ACR Vo. L5E17
efficiencies not only causes a decrease' In the maximum power obtainable but also considerably restricts the range of fuel-air ratio for which operation is possible. This fact is particularly evident in figure 1+. If the turbine efficiency is held constant and the compressor efficiency is varied, the power curves are quite similar to those shown. TA£ bU
1800
50
v
^
/
llj.00
,/*
\
/
\
-
-
\
R
15
30
s.\
\
?n C-\J
r N
12
85
i
\ \
600
(pei-eent'
V
\
1000
IT
\
-
l+o
16
i
8
65 R
'
1O
200
0
.OOlj.
.008
.012
.016
.020
.O2I4.
Wf/*a Figure I4..- Effect of fuel-air ratio on thrust and = pressure ratio. At sea level; V0 = 0; tmax 1500° P; = iTc 75 percent.
8
NACA ACR No. L5E17
2lj.00
2000
l600
1200
.012 .016 Wf/ffa Figure 5»- Effect of fuel-air ratio on thrust power and pressure ratio. At aea level; V0 = 800 feet per second; tmax = 15OO0 P.
NACA ACR No. L5E17
P
P/foa
900 - •
80
800 -
70
-1 -JI.
""
f
( percent.;
/" /
/•"**
.
\/
700 -
X
60
1
12
^J '
8 \
./ \
R
6;
600 -
50
^
\5
500 -
ko
.ooij.
.008
.012
.016
.020
.021;
Wf/Va Figure 6.- Effect of fuel-air ratio on thrust power and pressure ratio. Altitude, 1|.0,G00 feet; V0 = 880 feet per second; t^x = 1500° F; r\c = 75 percent. The maximum points of a number of curves of the type shown in figures I; to 6 are plotted on coordinates of oompressor and turbine efficiencies in figure 7 to show the relative importance of these two efficiencies, in this figure the axes may be interchanged with little change in the thrust or power curves, which indicates that, for all practical purposes when reasonable efficiencies are used, the thrust is equally sensitive to changes in turbine and compressor efficiencies and that the product of turbine and compressor efficiencies is more significant..than either efficiency alone. The effects of fuel-air ratio, pressure ratio, and maximum temperature on thrust and thrust power are shown in figures 8 and 9. Figure 8 shows the variation of static thrust at sea level with fuel-air ratio at each of three maximum temperatures. Lines of constant pressure
KACA ACR No. L5E17
10
90r \ -p
\
80
\v
\
\>
s
u
\
© Pi
^1600 \ \
VV
\ V
,.._
6o 6o
80
v
\
8o
\
\
P (Up)
W
\
ü
Pi
100
r—
\
SQ00
N
\ O
90
TVp, percent At sea level? VQ = 0.
(a)
a
1
I
70
90
]4oo
\
\" v \ Soo 100C L1200
L
-p
s.
\
70
T (115)
\
N>
«k
c>
V
V s. \
^
\
O
«=•
VN
70
\ \,
\ \,'
v >
65c
6o 60
1 1
70
V
•>
850 Vv
^
1 1
^8oo
's
1 7 oc !
80
750
90
100
T}p, percent (b) Altitude, lj.0,000 feet; V 0 = 880 feet per second. on thrust figure 7i- Relative effects of T)c and and thrust power, t.'max = 1^(;ooo P.
11
NAOA ACR No. L5E17
TA 2800
_ 90 R = 9 ....
2600
" 80
'
2ij.00
> <• /
/
c^ (y
2200 - 70
'\
/ /; '/ r
2000 60 l800
s
' /
-
c
/;
^ ''
>
' 2100 /
1t
max
S 1800
1500
1600 " 50
.c)10
•5
.0 ik
.0 18
.c 22
.026
Figure 8.- Effects of fuel-air ratio, pressure ratio, and = 0; maximum temperature on thrust. At sea level; T]c = TVp = 85 percent, ratio are drawn through these thrust curves. Figure 9 is a similar plot of thrust power at high speed and high altitude. These figures indicate that, for the range of temperature considered, the maximum thrust at any altitude is nearly proportional to the difference between free-stream and combustion-chamber temperatures. At any particular temperature the thrust is more sensitive to changes in fuel-air ratio than to changes in pressure ratio. Conroarisons of many curves of the type shown in figures I4. to 9 indicate that the pressure ratio and the fuel-air ratio for a particular power condition are primarily functions of the combined turbine and compressor efficiencies and the combustion-chamber temperature. Figure 10., which shows this relationship for the maximumpower oondition, is reasonably accurate for the range of flight speed and altitude considered in the present analysis,
12
NAOA ACR No. L5E1?
1 i) 10 R --» 11
130 600
l
-
* /
N v
/
110 500
'21UC
//
-
ij
100
'
V
hK
—
max
"P;\-l / /*v On
120 550
6
/
/ r
/1S00
,'
1
too
'// /
90
J
J4.OO /
60 350
I
l\
>J\
s.1500
/
/
70
300
60
250
— 50 .010
.Oil;
.01S Wf/Wa
.022
.026
Figure 9.- Effects of fuel-air ratio, pressure ratio, and maximum temperature on thrust power. Altitude, 60,000 feet; VQ = 880 feet per second; T]c = Tty = 85 percent.
13
NACA ACR No. L5E17
.0214.
^c'TT = 0 .65
'
.023
A
nc-rvp = 0.50
/
..022
WfAä
/
/ /
.021
t
.020 /
/ /
.019 Wf/Wt
/
.018 .017 .016
Y
/
/
>/
12
^' 'S
>1\
.015
,—*"J /
.Oll|. /
^"'
s*
10 8 6 It
.013
1500 l800 2100 Maximum temperature, °P Figure 10.- Effect of engine temperature on fuel-air ratio and compression ratio. Maximum-power condition,
Figure 11 shows how the thrust power and the thermal, Jet, and over-all efficiencies for 880 feet per second at J.0,000 feet vary with the fuel-air ratio. Turbine and compressor effioiencie., are held constant at 85 percent,
NACA ACR No. L5E17
ih
and the pressure ratio is varied to keep a maximum temperature of 1500° P or 2100° P. These curves show that the maximum-thrust condition is not the condition of most economical operation. If an engine having a maximum temperature of 2100° P (fig. 11(b)) is designed to run at maximum over-all ef "iciency instead of maximum thrust power, the Jet efficiency is improved from i+2 to 51 percent, the thermal efficiency, is slightly improved, andtho over-all efficiency increases from 21 to over 26 percent. The thrust power, however, drops to 1000 horsepower, only 70 percent of the maximum of lij.00. P/fc.
(percent) •
90
100
50
80
ko
60
30
ko
20
20
10
1000 >- "V \ /
900- 80 \
X
800- 70
\
/
\
/
/ ,f
^ \
f
*>J
\
700-
>r>
60
1 /
60050
/ /
5000 J+.OOI+.
Thrust power
\
\ \
Ü 1
.008 (&)
:^C
\
*t R
.012
.016
.020
W = 1500° P.
Figure 11.- Changes in thermal, jet, and over-all efficiencies and thrust power with fuel-air ratio. Altitude, lj.0,000 feet; V0 = 830 feet per second; T)c = T>p = 85 percent.
NACA ACR No. L5E17
15
lj.0
20
0
0
600 -
500•
li.0
.OOlj.
.008
. .012 .
.016
.020
••Vf/tfa
(*) Figure 11.- Concluded.
tmax =
2
100o P.
,02lj.
.028
16
< v
NACA ACR No. L5E17
' Figure 12 shows how the thrust power and the thermal, Jet, and over-all efficiencies vary-with the maximum temperature for operation both at maximum thrust power and at maximum over-all efficiency. As is to be expected, the thermal efficiency increases with the temperature range of the cycle and the jet efficiency decreases with the higher velocities that accompany the high temperatures. ?Jhen operation is at maximum power, an increase in maximum temperature does not cause a significant change in the over-all efficiency but the rats oi fuel consumption is considerably increased. For the maximum-power condition, therefore, the net result of using a higher ongine temperature is to increase the power capacity of the engine and thus to improve the power-weight ratio. When conditions of maximum over-all efficiency are specified, higher engine temperatures load to improvement in over-all efficiency as well as in engine capacity. Calculations for other flight speeds and turbine and compressor efficiencies show that, at maximum power, the over-all efficiency i3 nearly independent of maximum engine temperatui-e. The altitude effect is relatively small. Dhder these circumstances, curves showing over-all efficiency as a function of flight speed and the product of turbine and compressor efficiencies (fig. 13) will be approximately correct over the entire range of engine temoerature and altitude under consideration. In nearly all cases the over-all efficiency will fall within the ranges indicated in the following table: efficiency Product of turbine and Over-all (percent) compressor efficiencies ki\.0 fps 880 fps
0.1;
8 to 10
•5 .6
13 to 15
k. to 6
16 to 18
8 to 10
•7
20 to 22
10 to 12
Figures li; and 15 show how the thrust power per (pound per second) of air flow varies with flight speed and altitude. Figure li; describes operation at maximum
NAOA ACR No. L5E17
17
-
1I4.OO
-
1300
-
1200
-
1100
1000
1500
l800
-
900
-
800
2100
'max Figure 12.- Effect of temperature on power and efficiencies, Altitude, lj.0,000 feet; VQ = 880 feet per second; T)c = nrp = 85 percent.
18
NACA ACR No. L5E17
power and figure 15, operabion at maximum over-all efficiency. The fact that the curves for altitudes of 50,000 and 60,000 feet are coincident (fig. ll\.) indicates that the atmospheric temperature is the.only altitude effect which has a direct hearing upon the power per (pound per second) of conducted air. Z0 • +3
s
,880 fps
h
^kkO fps 10
'* j
/
0
IQ
.1
• +0
.(50
.80
1.
Figure 13»- Effect of turbine and compressor efficiencies and flight spoed on over-all efficiency.
PERFORMANCE OP JET-ETOINE INSTALLATION The turbine and compressor for a jet engine of the type under consideration should he so selected that the compressor torque required and the turbine torque produced exactly balance at tho desired rotational speed. The blade angles must be such that both turbine and compressor operate at maximum efficiency when the design air flow is obtained. An engine that has been designed for a particular maximum temperature, air-flow rate, and power condition (for example, maximum thrust and maximum over-all
NACA ACR No. L5E17
110
1 —
19
•—*.
1
L800 •,
/,
— 1500 0 p
1/
100
f
Al titude (ft)
f6o,ooo
//
"(50,000 / / //
90
1 Vl 7A /
8o P/V, 70
//
60,000 50,000•
6o
50
1/ Vl
AHbitu ie (ft)
itpjOCO25,000
Sea level •
e—lj.0,000 \ / //
7 Hi JII 1/ Iß If
/
/
'7 t 7lit
' / /
r>
/'Sea level
J
tf '// r
n // /
/ 25,000 /
/ /
/
/
k° /
200
I).00
600
800
1000
1200
Plight speed, fps Figure lir.- Engine performance when design is for maximum power. TJC = rm = 85 percent.
20
NACA ACR No. L5E17
Altitude (ft)
80
lj.0,000 f
25,000 70 , Sea level 60
PA« 50
to 7n
200
1|.00
600
800
1000
1200
Plight speed, ±"ps Figure 15.- Engine performance when design is for maximum over-all efficiency, t^x = l800° Pj TJc = T)rp = 85 percent. efficiency) may "be operated at the same temperature and power condition over a wide range of speed and altitude with little change in turbine and compressor efficiencies if the air-flow rate can be controlled in flight by an adjustable exit nozzle or similar moans. This simple adaptability of the jet engine js discussed in appendix B. Each combination of power condition and maximum temperature, then, actually represents a single engine, the performance of which follows with only slight discrepancies the foregoing calculations. The performance calculated herein is for an airplane powered with one of the jet engines described in the preceding section.
NAGA AGR No. L5E17
21
Engine,"- The turbine and-compressor operate, at adiabatic efficiencies of 85 percent, and a temperature of 15OO0 p is maintained in the combustion chamber. Operation is at the power, condition of maximum thrust per (pound per second) of conducted air. The exit nozzle Is of such a nature that the air-flow rate agrees with the curves In figure 2. The engine performance characteristics are shown by figure 16, in which thrust power is
3200t
ft
2J4.OO
ft
o p.
l600
-p
05
E-i
800
200
lj.00 600 8 00 Plight speed, fps
1000
1200
Figure 16.- Engine power output, tmax = 1500° P; T}c = Typ = 85 percent. plotted as a function of flight speed. At each altitude the thrust power increases almost linearly with the flight speed, which Indicates that the thrust force changes only slightly with flight speed. Airplane.- The airplane is a small, high-speed fighter-type airplane, with a gross weight of 5IJ.OO pounds and a wing loading of 50 pounds per square foot. The
22
HACA ACR No. L5S17
lift-drag ratio for tbe airplane is calculated by the equation CT
L D 0
errA
where oDo = O.Oli* A = 5-75 e = 0.9 For high flight Mach numbers this rr.tio is divided by the correction factor for C^ shown in figure 17, 3.2
I
•
f-:
O
2.4 •
•P
I
C
a
ß o O
/
1.6
>H
/
ti O
o
p
j .8
i
.8
1.2
M Figure I?.- ?Tach number correction for drag coefficient, (Fr 031 unpub 1 i shed data.)
MCA ACR No. L5E17
23
whi-ehwas obtained -f pom -unpublished., data. ..„ The. re suiting airplane power requirements for level flight are shown in figure l8.
3200
Pi
S 21^00 o a l600 •p
2
I4.OO
600
800
1000
Flight speed, fps Figure 16.- Power required by airplane for level flight,
Calculations of performance.- In order to calculate speed and rate of climb, the power required by the airplane for level flight at a particular altitude (fig. l8) can be plotted on the same graph with the power output of the engine at the same altitude (fig. lb). The intersection of the two curves will indicate the level-flight speed for the particular altitude-engine-airplane combination under consideration. For flight speeds less than that obtained in level flight, the excess of power available over power required may be used for climb.
HACA AGR No. L5E17
2fc
The variation in airplane level-flight speed with altitude is shown in figure 19• The fact that the speed
1000
800 m ft
6oo
I4.00
•6 &
200
w
0
20
T+Ö
"Sol' io5
Altitude, ft Figure 19.- Effect of altitude on level-flight speed of jetpropelled airplane, tj^^ = 1^00° F; T]c=T)p=85 percent; W = 5I4.OO pounds; S = lOo square feet. does not increase with altitude is due to the assumed effect of the Mach number on the drag. As the altitude increases, sonic velocity decreases and the Mach number effect on the drag is evident at lower flight speeds. If the airplane were 1 .rger or the engine smaller, the flight speeds would be lower but would increase slightly with altitude in the usual manner.
NACA ACR No.. L5E17
25
• ' In figure 20 the•maximum rate.of.climb is shown as a function of the altitude. The discontinuity in the slope of this curve occurs at. the altitude above whioh the temperature is assumed to be constant.
I
6000
5000 •H •P
Jj.000
S On
3000
-P cd
s
2000
1000 \
20
lj.0
60 x io5
Altitude, ft Figure 20.- Effect of altitude on maximum rate of climb of jet-propelled airplane. max = 1500°.p; TIC = Tjrp = 85 percent.
26
NAC'A ACR No. L5E17
These performances are only for a particular combination of airplane and jot engine. Curves of thrust power against speed for various altitudes may be drawn for an engine having any combination of turbine and compressor efficiencies and maximum temperature, and on these engine curves nay be superimposed the performance curves of any airplane. For an engine of different size, having flow properties similar to those of the engine described, the fuel consumption and thrust power vary with the square of the engine diameter. CONCLUDING REMARKS For each assumed maximum engine temperature in this performance analysis of jet-propulsion systems, particular attention has been given to only two types of cycle: the type giving moxiian . thrust power per (pound per second) of conducted air and the type giving the highest over-all efficiency of conversion of fuel energy into thrust power. When the turbine and the compressor are selected for engine operation at maximum over-all efficiency, the over-all efficiency is about li times and the thrust power 5 is about three-fourths of the corresponding values when operation is at maximum thrust power. The maximum thrust-power output at a particular altitude is approximately proportional to both the flight speed ard the temperature difference between the free stream and the combustion chamber. The efficiencies of the turbine and the compressor are about equally important in determining the engine performance. For reasonable values of each, the product of these efficiencies may be considered a good index of the attainable performance. The following table shows the relation between the product of turbine and compressor efficiencies and the over-all efficiency for two flight speeds. Although based only on cycles giving maximum thrust, the table applies to all altitudes, and engine -temperatures considered herein.
NAGA ACR Nö. L5E17
27
•Product of turbine and compressor 9fficleno les
"'Over-all efficiency (percent) IjJ+O fps 880 fps
0.1).
8 to 10
•5 .6
13 to 15
4 to 6
16 to 18
8 to 10
•7
20 to 22
10 to 12
The fuel-air ratios and the pressure ratios for the maximum-power cycles are determined by the maximum allowable gas temperature and the product of turbine and compressor efficiencies. The approximate magnitudes of these ratios for three gas temperatures are shown In' the following table: ITaximum temperature
Pressure ratio
1500
•j to 6
1800
8 to 9
.019 to
.016
2100
10 to 12
.022 to
.020 "
f°P)
Fuel-air ratio 0.015 to 0.013
In this table the prodvict of turbine and compressor efficiencies Is constant at O.65. A similar table for an efficiency product of 0.50 would show fuel-air ratios 0.001 to 0.002 higher than and pressure ratios about two-thirds of those shown. Langley Memorial Aeronautical Laboratory National Advisory Committee for Aeronautics Langley Field, Va.
28
NACA ACR No. L5E17 APPENDIX A SYMBOLS
a
cross-aectlonal area of duct, sq ft
A
aspect ratio
o
sonic velocity, fps
CD
drag coefficient
Gj)0
profile-drag coefficient
Cj,
lift coefficient
d
diameter, ft
D
drag, lb
e
span efficiency factor
L
lift, lb
M
Mach number
n
rotational speed, rps
p
pressure, lb/sq in.
P
thrust power, hp
3
voliime rate of flow of air, cu ft/sec
R
ratio of static pressure after compressor to static pressure before compressor
S
wing area, sq ft
t
temperature, °P
tjjL^
maximum gas temperature, °P
T
thrust, lb
v
specific volume, cu ft/lb
V
velocity of air through duct, fps
(V0/c)
NACA ACR No; TJyElJ
2?
V-
velookty of-air through exit nozzle, fpa
VQ
flight speed, fpa
W
gross «eight of airplane, lb
Wa
weight rate of charge-air flow, lb/sec
Wf
rate of fuel consumption, lb/seo
n
adlabatio efficiency of meohanioal compression, ratio of isentropic to actual enthalpy increase for a particular pressure -rise
nn
thermal, or cycle, efficiency (Heat input - Heat reJeotedN Heat input J
TJT
propulsive efficiency of the air jet
Tit
over-all efficiency of conversion of fuel energy into thrust power
T)
adiabatic efficiency of turbine, ratio of actual to isentropic enthalpy decrease for a particular pressure drop
p
air density, lb/cu ft
J
/ 2VQ \ f *—\
V'o
+
V
30
NÄCA ACR No. L5E17
APPENDIX B CONDITIONS AND ASSUMPTIONS The conditions and assumptions used in the foregoing analysis are listed he-ein. Some explanation is given when necessary. Army summer air is used exclusively. The thermalair Jet engine operates on the same cycle as a ducted cooling system; the atmosphere customarily used In calculations for cooling equipment Is therefore used for the Jet engine. This use of Army air gives a somewhat conservative estimate of engine performance. Uniform temporeture, pressure, and velocity exist over any cross section of air duct in the engine. Air velocity throughout the engine is kept so low that deviations from this idealized condition are of little significance. No heat is lost from the engine by conduction, and all air flow is frictionless except In the turbine and compressor. These idealizations Invalidate none of the conclusions, since the duct friction losses are negligible with the velocities used end the conduction heat losses may be taken care of by using a slightly higher fuel rate. It may be possible in some cases to Improve the performance of a Jet-engine installation by decreasing the duct area and therefore the frontal area of the engine and by taking some friction loss in the ducts. The weight flow of charge air Is maintained proportional to the stagnation density of the charge air. The following discuss? on indicates the reasons for this assumption and some of Its results. With this air-flow control, the velocity of the air entering the compressor is practically constant. Tfhen the engine Is operating at maximum thrust power and a particular maximum temperature, the velocity of air entering the turbine Is constant to about ±7 percent, for flight speeds of I4J4.O to 880 fps at altitudes from sea level to 60,000 ft. The continuity equation for the conducted air is pVa = W,a
31
NACA ACR No. L5E17
The entrance "area and velocity remain" cons taht" for "both compressor and turbine. The weights of the air flowing through the compressor and the turbine are essentially equal and thus vary at the same rate. The density, of the air entering the compressor and the turbine, therefore, varies at the same rate« Furthermore, both.the torque output of the turbine and the torque required to run the compressor at constant speed are proportional to the density of the conducted air. The compressor torque required and the turbine torque produqed, therefore, vary at the same rate and remain balanced at a constant rotational speed over a wide range of altitude and flight speed. Because the assumed control of air flow makes it possible to operate the compressor at constant Q/nd^, the power put into a pound of air by the compressor Is constant for all altitudes and flight speeds. The pressure ratio is greater at high altitudes, then, than at low altitudes and Is slightly greater at low flight speeds than at high flight speeds. In fig. 21 the *o 12 -ifpa)-
0
*r
11 .•
10
k^
•p
cd
u
1
m»
r
66o R Rn-
'/ '/
9
& a m a>
Design point
10
20
30
ko
50
6o x io5
Altitude, ft Figure 21.- Effect of altitude on pressure ratio over an axial compressor. Constant rotational speed and air volume.
52
NACA ACR No. L5E17
pressure ratio is shown as a function of altitude and flight speed for an axial compressor operating at constant Q/nä? and an adiabatic efficiency of 85 percent. This variation in pressure ratio for a particular compressor causes an engine designed for maximum power at a particular altitude and speed to operate at maximum power over a wide range of altitude and speed. The heat value of the fuel is 19,700 Btu/lb, and combustion is assumed complete at the turbine entrance. Losses in total pressure occur only in the compressor and the turbine, and mechanical losses are allowed for in the efficiencies of the compressor and the turbine. Because of the great excess of air, the properties of the exhaust gas are assumed to be the same as those of air, but allowance is made for the increase in the weight of the conducted air due to the addition of fuel.
NACA AGR NO. L5E17
33
REFERENCES
1. Salisbury, 0. Kenneth: The Basio Gas-Turbine Plant and Some of Its Variants. Oil & Gas Jour. v Pt. I, May 13, ]9Ul-» PP. 59» 62, 63, 66, and Pt. II, May 25, I9I44, pp. 101, 102, IO5, 107. 2. Lutz, 0., and Wolf, P.: IS-Tafel für Luft und 7erbrennungsgas9. Julius Springer (Berlin), 1938. 3. Palmer, Carl B., and Brevoort, Maurioe J.: of Cooling-Air Flowers on Thrust Power. No. li±G2lj., I9I+I+.
The Effect NACA ARR
> o > > o o r m
jp
./e
?o
Entropy, Btu/C/b)CF obs) Figure /. - Thermodynamic properties of air, (Transformed from reference 2)
OR
Palmer, Carl B.
lAITO-
782U
DIVISION: Power Plants, Jet and Turbine (5) ORIG. AGENCY SECTION, Performance (16) CROSS «EFEaENCES, Engines, Turbo-jet - Performance data ACR-L5E17
NUM&EB.
CA155)-
AUTHOa(S)
AMER TITLE, performance of compressor-turbine jet-propulsion system fORG'N. TITLE: ORIGINATING AGENCY,
National Advisory Committee for Aeronautics, Washington, D. C.
TRANSLATION, COUNTRY LANGUAGE |FOKG'N.ClASS U. SXLASS. Eng. U.S. Unclass.
DATE
PAGES
Jun'lt? I $k
11LUS.
27
FEATURES
tables, graphs
ABSTRACT
Analysis of performance was carried out by calculating the thrust power from compressorturbine jet engine with a systematic variation of pressure ratio, compressor and turbine efficiencies, flight speed, etc. Increasing compressor and turbine efficiencies from 70 to LO percent doubled over-all efficiency of engine at J00 mph. Speed of 300 to 600 mph increased over-all efficiency from 7 to 10 percent. Maximum output at particular altitude was approximately proportional to temperature difference of combustion chamber and free atmosphere. fiOTEi Requests for copies of this report must be addressed to: N.A.C.A., Y.'ashington, D. C. T-2, KP. Alu MATERIEL COMMAND
Alu
VECHNICAL ONDEX
WSIGHT FIELD. OHIO. USAAF if} Bf-O-31 OAfl «> IK,
TITLE: Performance of Compressor-Turbine Jet-Propulsion Systems AUTHORS): Palmer, Carl B. ORISINATIN6 AGENCY: National Advisory Committee for Aeronautics, Washington, D. C. PUBLISHED BY: (Same)
June '45
Unclass.
U.S.
ABSTRACT:
Eng.
34
tables, graphs
over An analysis of the performance of compressor-turbine Jet-propuls ion systems was carried out by calculating thrust power from a compressor turbine Jet engine with a systematic variation of pressure ratio» fuel-air ratio, compressor and turbine efficiencies, flight speed, altitude, and maximum gas temperature. It was found that by increasing turbine and compressor efficiencies from 70 to 80%, over-all efficiency of the engine was doubled at 300 mph. Maximum power output at a particular altitude was found to be approximately proportional to temperature difference between combustion chamber and free atmosphere.
DISTRIBUTION: Request copies of this report only from Originating Agency DIVISION: Power Plants, Jet and Turbine (S) SUBJECT HEADINGS: Engines, Turbo-Jet - Performance data SECTION: Performance (16) (34155); Engines - Performance calculation (32874.2) ATI SHEET NO.: R-5-16-21 ' Document* Div.iion IntaMigance Department Air Materiel Commond
Wright-PaHenaa Air Fort» taM
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