ASSESSMENT OF TURBOJET ENGINES

Chapter 4. An appreciation of subsonic engine technology

4.4. ASSESSMENT OF TURBOJET ENGINES

analysis of the engine, taking various en-gine des·ign parameters as variables. En-gine manufacturers have computer programs for performing these studies, and the same applies to some aircraft companies. Meth-ods are also presented in the relevant lit-erature (Refs. 4-25, 4-26, 4-29 and 4-50).

In Appendix H a summary is given of ana-lytical expressions which, in spite of the simplifications introduced, are suitable for making an initial estimate of engine performance in preliminary aircraft design.

This method has been used for the gener-alized calcul"ations which follow. When en-gine performance is calculated a distinc-tion can be made between:

1. performance at the design point, 2. off-design performance.

Design point is intended to indicate the working condition at high rpm where the efficiencies of the compressor and turbine are optimum. The design point may be chosen for a representative situation, such as the takeoff, climb or cruise. In this way the engine can be adapted as far as possible to its use in the aircraft to match the most critical performance requirement. In the case of the straight jet engine, the design point is generally assumed as the takeoff, whereas the most critical condi-tion for high bypass engines will often be the cruising flight.

mainte-:x:l1.5

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z 0 i= ~ 1.3

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z 0

u12 -' w :::l u..

u

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r.Jl ~ 0 ~ 1.0 u w cr: cr:

8

^.9

.8 .7

1200 K 3

.6

OPTIMUM FAN PRESSURE RATIO

M. :.85

NOTE: This diagram is drawn for the tropopause, but is equally valid for other altitudes, provided T is

t4 chosen so that the same value of

·5o!:---=1o!:;---;;2!;:;0----:3obo,--4.-J;o,--s;::~;o,.---7faon--. -::;7'0

CORRECTED SPECIFIC THRUST }L?s-SEC

Tt /T is obtained 4 0

Fig. 4-22. Effect of OPR and TET at cruising conditions

nance cost increase with complexity.

4.4.1. Overall Pressure Ratio

When a cycle analysis is carried out on turbofan engines with varying working pres-sures and temperatures at the design con-dition, a convenient representation of the results can be given in the form of Figs.

4-22 and 4-23. The corrected specific thrust and fuel consumption have been used in order to make the result valid for dif-ferent altitudes. This figure shows that in all configurations an increase in OPR leads to a reduction in the specific fuel consumption in cruising flight.

In order to obtain a high specific thrust,

either in cruising flight (Fig. 4-22) or during takeoff (Fig. 4-23), a high value for the OPR is not required at the given TET. In the case of the pure jet engine, for example, with TET =1,000 to 1,200K in cruising flight, the condition for maximum specific thrust is ec = 7 to 9 and in the takeoff between 9 and 11. In the case of turbofan engines these values will be slightly higher, say 12 to 16 at the gen-erally acceptable turbine temperatures

(Section 4.4.2). In view of the fact that - within certain limits - an increase of OPR over these optimum values has little influence on the specific thrust, the de-signer generally aims at higher pressure ratios which lead to a better fuel

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0 PROGNOSIS·

"=

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Tt4"" 1800 K

.2 ^{Ec= 36}

20 30 40 50 60 7 80 . 90

SPECIFIC THRUST T/W -SEC

Fig. 4-23. Turbofan engine takeoff performance

50 0 i=

~40 w a: ::J (/)

~30

...

-'

~

²⁰

~

'55 '60

0 AXIAL HOW 0

$ A~IAL•CENHII / / --0 --:-.::;-...,

~/ eCENTRIFUGAL

! ..._---~L'L' ^ENGINES

'65 '70 '75 '80 '85 YEAR OF CERTIFICATION

Fig. 4-24. Historic development of the OPR (References: Jane's All the World's Air-craft, various engine manufacturers' data)

sumption in the design point and also at reduced rpm. Up to about 1953 jet engines were designed with a single stage centrif-ugal compressor (Fig. 4-15) which gave a maximum OPR of around four. As these en-gines had a high fuel consumption and large frontal area, they have been superseded by engines with an axial compressor (Figs.

Fig. 4-25. Specific fuel consumption of straight jet engines at sea level

4-15 and 4-19a). There are some recent ap-plications of combined axial and centrif-ugal compressors on relatively small en-gines, such as the Garret AiResearch ATF-3, which have quite acceptable pressure ratios and a small frontal area.

Axial compressors make it possible to ob-tain large pressure ratios and a small

frontal area. A value of £c ~ 8.5 is still possible with a single compressor and sin-gle spool. When this value is exceeded, the design conditions for the foremost and rearmost compressor stages are so far apart, that it will be necessary to employ two compressors, running at different rpm: twin spool engines.

The highest value for the OPR actually used in straight jet engines is about thirteen.

Fig. 4-24 shows how the pressures have risen during the fifties. Fig. 4-25 gives an impression of the influence of OPR on CT during takeoff. The actual fuel consump-tion of a number of engines is compared with calculated values.

The tendency towards high pressure ratios has been continued in the turbofan engine up to values of 25 to 30 in the case of the latest generation for large transport air-craft. Fig. 4-26 gives a schematic

repre-A. Twin spool; fan com- 8. Twin spool; separate bined with low pressure fan

section of compressor

C. Triple spool D. Twin spool; geared fan

Fig. 4-26. Principle of various fan-com-pressor configurations of high bypass en-gines

sentation of some engine configurations in which these high pressure levels can be a-chieved.

Layout A has the disadvantage that the last stages of the low pressure compressor con-tribute little to the increase in pressure, as they operate at relatively low tip speeds.

Layout B does not have that disadvantage, but demands a very high pressure ratio from the high pressure part, which entails the use of variable incidence guide vanes.

Layout C, employing three shafts, is at-tractive for various applications, since a relatively small pressure ratio is de-manded from each of the compressors (Ref.

4-34}. The specific weight is lower than that of the others.

Layout D, the two-shaft engine with geared fan (see also Fig. 4-19d} appears to be particularly suitable for relatively small engines where the power to be transmitted by the reduction gear is limited. The re-duction results in a low fan tip speed, while the turbine is still working at a

relatively high tip speed, resulting in a compact turbine. For several existing or projected geared fan engines thP gear ra-tio amounts to between .25 and .50.

From Fig. 4-22 it is seen that for an en-gine with A ~ 8, for example, a pressure ratio increase of up to 36 will only give a modest gain in the specific fuel con-sumption. At these high pressures the tem-perature at the rearmost compressor blades becomes so high that the material proper-ties will deteriorate. For a future gener-ation of large engines for long-range transport aircraft £c ~ 36 may possibly prove to be the limit (Fig. 4-24}, but there is also a notable tendency towards considerably lower values in order to a-chieve simplicity coupled with low cost.

In the future we may expect a range of values between Ec ~ 12 and 40, depending upon the use for which the engine is de-signed. The author's forecast is as fol-lows:

- small executive aircraft, feederliners:

12 to 14

- short-range transport aircraft: 20 to 25 - long-range transport aircraft: 35 to 40

4.4.2. Turbine Entry Temperature

Returning to Fig. 4-22, the following may be observed:

a. Ror a given bypass ratio a TET value which gives a minimum CT can be found for every OPR. This is henceforth indicated as the "optimum" TET.

b. The optimum TET increases with increas-ing bypass ratio, the reason beincreas-ing that for high ^~ the propulsive efficiency is not so sensitive to increasing TET as it is in the case of low bypass engines (Fig.

4-27).

2000.---~3600

"

₁₁₈₀₀

~ BOUNOAAY OF ACTUAL T14

i:1600

ffi ..

~1400

ACTUAL ENGINES CRUISE CONDtTIONS

a: I

2800

2400

OPTIMUM TURB.ENT.TEMP.Tt4 AT 2000

CONSTANT OPTIMUM 0\ffRALL

~s~~~~t~~L~ON;.⁵ 1600 8000~---2~---4~--~6----~8----~10--~12

BYPASS RATIO

Fig. 4-27. Actual and optimum TET in cruis-ing flight (civil engines)

c. At high bypass ratios, TET may be varied within fairly wide limits above or below the optimum value without greatly affecting CT. At ~ : 8 and Ec : 28, the TET will be optimum at 1500 K (2/00 R) in the example presented. However, when the TET is chosen 10% lower (1350 K, 2430 R) it will cost only about 1% in CT.

d. For all combinations considered, a high TET is favorable for obtaining a high spe-cific thrust.

The maximum TET values used in actual en-gines for cruising conditions (Fig. 4-27) are generally higher than the "optimum"

values. On the one hand, a high specific thrust has the advantage of low engine weight and installed drag, while, on the other, it will mean that in normal long-range cruise the engine will not operate at the maximum permissible cruise TET, but at a lower value which is closer to the

op-timum for low fuel consumption and long operating life.

Maximum values for TET are higher during takeoff than in cruise (Fig. 4-28) as they

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1-·r---~3~

0

rr: I 3000

2~

2000 800·~--~--~--~~--~--~~--~--~~~

1950 '55 '6o '65 '70 '75 'so '85 YEAR OF CERTIFICATION

Fig. 4-28. Historic development of maximum TET during takeoff (References: engineman-ufacturers' data)

will generally be required only for a lim-ited period of time. In the case of low bypass engines, the difference will be in the region of 150 to 200 K (270 to 360 R), but they will be less (50 to 100 K, 90 to 180 R) for high bypass ratios, since the size of the engine is determined by the thrust required in cruising flight.

At present, with the use of uncooled tur-bine blades, temperatures of about 1200 K

(1260 R) may be regarded as permissible for short periods. When they are air-cooled temperatures of up to 1600 to 1650 K (2880 to 2970 R) may be possible. Still higher values may be achieved in future with the use of transpiration cooling'and improved materials, although there is a considerable difference of opinion among experts on the actual gains to be expected.

Looking at the \Talues used in practice over the past years (Fig. >-28), it may be noted that the average increase comes to about 20 K (36 R) per year. If this trend con-tinues in the future we may expect a gen-eration of large, high bypass engines in the eighties with a TET of 1800 K (3240 R) during takeoff and 1750 K (3150 R) duriny cruise. The data in Ref. 4-48 indicate that these temperatures are feasible, with only moderate quantities of air required for cooling.

4.4.3. Bypass ratio

The choice of this parameter has such. a far reaching effect on the design of the engine and its installatioh, that the optimization of A cannot be fully dealt with here, in view of the complicated nature of the prob-lem. Accordingly, we shall confine our-selves to summarizing some of the more im-portant aspects.

a. Design of the fan and low pressure tur-bine.

The power supplied by the gas generator may be divided in various ways over the hot and cold flows. The division ^~f power can be characterized by the work ratio,

work ratio= power tran~ferred to fan ( 4_29 )

convert~ble energy This ratio is controlled by the pressure ratio of the low pressure turbine and the Fan Pressure Ratio and determines the ex-haust velocities imparted to the hot and cold gases. It can be shown (Ref. 4-30) that the propulsive (and overall) effi-ciency will be maximum when:

(.70 to .75 are

typical values) (4-30)

where nt and nf are the turbine and fan efficiencies respectively. Fig. 4-29 shows

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that for moderate values of A, the Fan Pressure Ratio (FPR) has no great influ-ence on the specific fuel consumption, at any rate within certain limits. If, say, for A = 5 we take an FPR of 1. 55 instead of the thermodynamic optimum of 1. 66, this will result in an increase of only 1\ in CT. The variation in the power absorbed by the fan, however, is closely related to that of the turbine. In view of the lim-ited pressure ratio per stage, the number of turbine stages will have to be adapted for certain combinations of A and FPR. If we now vary the FPR with a given number of turbine stages, the optimum will be found to lie at another value than in the pre-vious case.

In Fi~. 4-30 the optimum FPR is given as a

OPTIMUM VALUES. CALCli.ATED FOR:

M0 • .85

h,.10,000m (32,800 FT)

1.5 t;,

36 14SO K 12610 Rl 26 1100 Kl1980 Rl

,.

1.00 2 4 6 10 12 14 16

BYPASS RATIO

Fig. 4-30. Calculated and actual FPR function of A for some combinations ofOPR and TET, calculated for M = .80 with the analytical method given in Appendix H. The points plotted in this figure represent values for the takeoff applicable to ac-tual engines; co~responding values for cruising flight are higher (Section 4.4.2).

The graph shows that for a given gas gen-erator cycle the optimum value for the FPR decreases with increasing A. For high by-pass engines this effect will be partly compensated by the higher temperature and pressure levels which will generally be chosen in such cases, resulting in agreat-er specific convagreat-ertible powagreat-er. The maximum Fig. 4-29. Effect of fan pressure ratio on value for the FPR with a single stage fan SFC and number of low-pressure turbine stages will be about 1.60 to 1.70.

When the bypass ratio is increased at a given rpm, the fan diameter and tip speed will also increase. In the case of many modern ~ngines, part of the fan will work at transonic speeds which, amongst other things, will cause an increase in noise level (sub-para. f). This may be improved by using three engine shafts, thus reduc-ing the rpm of the fan, or reducreduc-ing fan speed by means of a reduction gear (geared fan, see Fig. 4-19d).

The Oowty-Rotol Company has developed a geared fan with rotor blades which can be adjusted in flight, so that an optimum ad-justment may be obtained for various en-gine settings. Thrust reversal can be a-chieved by giving the rotor blades a neg-ative pitch. The extra weight of the blade adjustment amounts to about 8% of the en-gine weight, whereas it will be over 15 to 20% in the case of a conventional thrust reverser for ~n engine with a high bypass ratio (Refs. 4-47 and 4-52)

b. Performance in high speed, high alti-tude flight.

The influence of A on the specific thrust and fuel consumption for a number of com-binations of OPR and TET is given in Figs.

4-31 and 4-32. The following may be noted:

~~~1.3

BYPASS RATIO

u y;

hl ^.9

..

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@ _RATIO

.... _{. 8} hl a:

a: 0 u .7

~-·oPTIMuM FAN-PAEsSUAERATIO:

, M0 ., .85 '

: Tt .1500 K (2700 R) I i INTAKE PRESStRE LOSS 4% OF FREE

20 30 40 50 60 70

CORRECTED SPECIFIC THRUST ~-SEC .6 0

I STREAM DYNAMIC PRESSURE 10

Fig. 4-31. Turbofan performance in cruising flight as affected by OPR and bypass ratio at constant TET

1. In the case of engines without intake and exhaust nozzle losses, CT will continue to decrease with increasing A, provided the optimum FPR is chosen for each configura-tion. A theoretical minimum will be reached with A + ~. corresponding to a propulsive efficiency ntf = nt.nf.

2. The specific thrust decreases rapidly with increasing A, although this effect may be partly compensated by choosing higher values for the TET.

3. When intake and exhaust nozzle losses are accounted for, there will be a bypass ratio for which CT is minimum. The effect of these losses is greatest when A is large and will even be accentuated under opera-tional conditions by the power extraction required to drive aircraft systems and bleed air to the de-icing and air condition-ing systems. The minimum installed CT which may be achieved, and the corresponding val-ue for A, are largely dependent on the mag-nitude of these losses (Fig. 4-33).

4. Fig. 4-32 shows three combinations (points A, B and C), which are representa-tive of different generations of engines, both recent and present, for transport air-craft. A radical improvement in CT has been achieved ln the present generation of large engines like the Rolls-Royce RB-211, Pratt and Whitney JT-90 and General Electric CF-6 high bypass ratio turbofan engines.

The development of future generations may take different courses.

Point o 1 : A bypass ratio of about 9 at a TET of about 1750 K (3150 R) and an OPR of about 36. The improvement in fuel consump-tion as compared with the current genera-tion is approximately 9% non-installed and 8% installed, while the specific thrust in cruising flight is about 15% less. In view of the high flying speeds of long-haul transport aircraft, a much larger value for A may lead to excessive engine diameters and installation drag and will probably not be favorable.

Point o 2 : Very large bypass ratios of, say, 20 to 30 at a TET of 1600 K (2880 R) and an OPR of 25. This configuration will imply a

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0.. CJ) NO LOSSES

WITH LOSSES

OPTIMUM FAN PRESSURE RATIO FOR MINIMUM CT

M0^~ .85

I: \-1050K<1890R>

Ec.12

ll: \•1250 K <2250 R>

£c.20

][. T14.1450 K <2610 R>

Ec=28 0 w

t-u

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a: 0

u

nz:: \=

1800 K <3240 R>

.5~----~---~----~---~---'~----~---~---£~c~=~3~6~~---~----~

0 10 20 30 40 50 60 70 80 90 100 110

CORRECTED SPECIFIC THRUST

~-SEC

Fig. 4-32. Effect of bypass ratio and intake/exhaust losses on performance in cruising flight

u

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~ 6

...

z

~4

a:

~ 2

M0 • • 8 11> 30.000 FT (9150 m I

\ • 1450 K(2610 R) Ec-24. >..5

0

o

1 2 3 4 5

BLEED·AIR FLOW, %OF GASGENERATOR FLOW Fig. 4-33. Effect of bleed air takeoff on SFC

very large fan diameter ("prop-fan") and entails the use of a reduction gear. Fuel consumption may be about 15% less at the flying speed chosen for the case in Fig.

4-32, but the gain will be greater as the cruising speed decreases. The drag of these engines increases rapidly with cruising speed.

The engines in this forecast are suitable

for different types of aircraft. Type o 1 may be particularly useful for fast long-range transport aircraft, while type o 2 may possibly be used on smaller and less rapid short-haul aircraft which should have good takeoff and landing performance and avery low noise level (sub-para. f) .

The effect of altitude on engine perform-ance is dependent on many factors. An ap-proximation for the thrust lapse with al-titude is occasionally found in the liter-ature, for example:

thrust at altitude

thrust at sea level an (n < 1) (4-30)

where both values of the thrust have been defined at the same Mach number and engine rating. Equation 4-30 should·be considered as an interpolation method for calculating ungine performance at altitudes where the engine manufacturer has not specified the thrust, rather than as a prediction method.

The designer should refer to Appendix H or employ any other suitable method available

127

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NO INTAKE LOSS

ti~

^Tt4 MAX. CRUI.SE VALUE

u 1.0 u. (/)

0 w .9 u 1-_w

a: a:

0 u .8

.7

.6

.50 2

when he needs a prediction of the thrust lapse with altitude.

Although engine performance can easily be predicted by using the method given in Ap-pendix H, the working pressures and tem-peratures of the engine should either be known or chosen. When the designer is prepared to accept statistical averages for cruising flight, reasonably accurate fig-ures can be found from Fig. 4-34 for effi-cient engines with high OPR. As against this, a comparison with Fig. 4-31 shows that for low OPR values the SFC will be considerably higher.

c. Takeoff performance and other flight conditions.

The specific fuel consumption is of less importance for performance in takeoff and climb, and we shall therefore pay rather more attention to the thrust lapse with speed, altitude and engine setting.

If Tto represents the static thrust at sea level, and if at the same time we assume that the gross thrust and mass flow through the engine do not change appreciably over a speed ranqe of up to about M

=

.15, we may write:

12R

6 7 8

T Tto

BYPASS RATIO

Fig. 4-34. Statistical values of SFC in cruising flight

(4-31)

In this expression T decreases linearly with vo due to the effect of ram drag

m

^vo.

At higher speeds, however, the effects of dynamic pressure on engine operation must be taken into account.

Introducing the specific thrust into (4-31) it follows that:

T Tto

34.714 _M0

1 - (4-32)

This equation may be refined by taking ac-count of the fact that the gross thrust in-creases with speed due to the dynamic pres-sure and that this is intensified as the bypass ratio increases. Assuming the mass flow through the engine to be constant, we may deduce (Appendix H) :

Here G is the gas generator function, de-fined in Section 4.3.1, which will typical-ly be of the order of .9 to 1.2 during takeoff.

From inspection of (4-32), it will be clear that since the specific thrust decreases

with the bypass ratio (Fig. 4-23), the takeoff thrust of a high bypass engine will decrease more rapidly with speed than is the case with a low bypass engine. Fig.

4-35, which is based on empirical data,

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a: :I:

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0 w .6

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~

.5 0 .05 .10 .15 . 20 .25 .30 MACH NUMBER

Fig. 4-35. Statistical-curves of thrust lapse during takeoff

may be used for performance calculations in preliminary design. It shows the vari-ation in thrust, deduced from engines with A ranging from 0 to 7. Both curves for A

=

9 and 12 apply to a TET of 1800 K (3240 Rl and an OPR of 36, and are computed with 4-33.

Some of the consequences of the above are as follows:

1. If, for a particular aircraft design, the thrust level is based on a requirement relating to the runway length or climb performance at low flying speed, the air-craft which is fitted with low bypass en-gines will have more thrust in cruising flight than one with high bypass engines, since the thrust decreases less rapidly with speed and altitude.

2. If two aircraft designs, which in other respects are entirely similar, have the same maximum cruising speed at a certain height, while their engines have the same installed cruise thrust, the high bypass engine will generally result in better takeoff and climb performance.

For given aircraft design requirements for cruise and takeoff, it is possible to find a bypass ratio where the thrust required in both conditions will be matched and this might form a basis for choosing A. In actual practice there are still other factors to be considered:

1. Cruising altitude is often a fairly arbitrarily chosen parameter which may still be varied within certain limits.

2. The engine manufacturer can influence the engine thrust lapse to some extent by adaptation of the compressor design point. When, for instance, the emphasis lies on the takeoff requirement, the de-sign point of the compressor is chosen accordingly so that the compressor efficiency will be maximum during takeoff, which implies that it will be lower during cruising flight. When, however, the cruising flight is taken as the sizing condition, the design point will be chosen with this requirement in mind • 3. The engine manufacturer will design the engine for use in several types of aircraft. Once the en-gine configuration has been chosen it is still possible to influence performance to a limited ex-tent - for example by increasing the TET, which will increase the quantity of air flowing through the gas generator, though the engine will work slightly less efficiently with regard to fuel con-sumption.

d. Engine weight.

From the decrease in the specific thrust with A (Fig. 4-23) one might tend to con-clude that both weight and drag due to the powerplant installation will increase. It will be seen from Fig. 4-36 that this

rea-!~~MR~.~.~-~~~M~~~;---,

~ ~.4

~

~

e

;:2

~-~-1

o0 o

0 OOo

0

15 jj,j:jO, ..

loL----L----~--~L----L----~--~~--~-·-·--~~

_{0 2 3 4 5 6 7 8}

BYPASS RATIO

Fig. 4-36. Specific dry weight of turbojet engines. (References Jane's All the World's

Aircraft, engine manufacturers' data)