Chapter 4. An appreciation of subsonic engine technology
4.5. ASSESSMENT OF TURBOPROP ENGINES
Contrary to expectations in the fifties, the development and use of turboprop en-gines in civil aviation has not been so widespread as that of turbojets. Tt.is is mainly due to:
a. the relatively great complexity of the co~ination of a gas, turbine engine with propeller reduction gear and constant speed propeller, as compared with the straight jet engine,
b. the higher flying speeds which can be attained with the jet engine, as compared with the propeller engine.
In the competition with the piston engine, the turboproo was only able to take the lead when technological progress enabled compressor pressure ratios to be raised to about 5 or 8 and Turbine Entry Temperatures of about 1200 to 1300 K (2160 to 2340 R) became feasible. This has resulted in the turboprop unit completely replacing the piston in the 5~0 to 3000 hp class of civil
'··
engines. Where engine first price is an important factor, however, the turboprop unit is still very much at a disadvantage.
4.5.1. Performance
The power developed by the gas generator is distributed partly to the propeller shaft and partly to the engine exhaust gases. It can be shown that the propulsive efficiency and the flying speed are the deciding factors in the optimum distribu-tion of energy. The following optimum jet efflux velocity is found (Ref. 4-7):
v 0 (4-39)
It is concluded that when the flight speed in the design conditions increases, the optimum gas efflux velocity will also in-crease. Consequently, when a turboprop en-gine is designed for relatively high speed operation, the thrust of the exhaust gases will be considerable at low speeds. If, on the contrary, the main emphasis is on high power output at low speeds, a low efflux velocity will be desirable (turboshaft en-gines).
~lr
"''i;.6
u z
Q
t:
.5~ (/)
8
4..J w
ii' u u: .3
8
Cl.(/)
.2
.1
50 100 150
2000 2200 2400 - R 14r---.---.---.---.---, 1800
COMPRESSOR CONFIGURATION:
o 1 CENTRIFUGAL 0 Cl> 2 CENTRIFUGAL
~ 12 • AXIAL
a: ® AXIAL +CENTRIFUGAL
~~ENGINE DEVELOPMENT w a:
::::>
~ w 10 a:
"-..J ..J
/ / / .
Y---o-~----7' ®<(
a: 8
~
(c FOR MAX. SPEC. / 1 ® POWER AT / I@) fl.--GIVEN 1: / I <1 ;.-.~6 !CRUISE I
\;Y
@@ ~~ ®II> llif/ @-'fY',®
~ I /_@
)-/ '
4 o-lV/ ~
2L---~----~----~----~--~
900 1000 1100 1200 1300 TURBINE INLET TEMPERATURE
1400 -K Fig. 4-43. 'l '.peratures and pressure ratios of turboprop engines (data from Ref. 4-27)
When (4-39) is applied to a family of tur-boprop engines, the specific fuel consump-tion and power may be computed with the method explained in Appendix H. This has been carried out for a cruising speed of
SHAFT HP
200 .LB SEC
M0 •. 40
H0 .6000m(19,680ftl
1. 4 ~i
1.2
1.0
.8
.6
.4
.2 Fig. 4-42. General-ized performance of
?oo
~----~----2-0~0~--~----~30~0~--~---4~0~0~--~----~5~0~CORRECTED SPECIFIC POWER ~-SHAFT HP W6 KG/SEC
turboprop engines at cruis11 1 condi-tions
~~
.6 1 AXIAL FLOW~~
uh. COMP . =¢1 1.2
z . 5
0 i=
a.. ::!:
~ .4
en ~ u
...J
w .3
~ LL
~ LL
u
w .2en a.. NOTE: Symbols as in Fig. 4-43
,4
.12 3 4 5 6 7 8 9 Fig. 4-44. Specific fuel
con-sumption at sea level static conditions
OVERALL PRESSURE RATIO
LB/SEC HP a...5r2r5----~--so~~---7rs ________ 1,o_o ________ 12Ts _________ 1so~~ a..
~
~"'
afl!+
~ l
SEA LEVEL/I SAJSTATIC TAKEOFF
1.0~
.8 1-~·3
w ~
!1;.2 w (!) z w
~ ~.1
0.. Vl
.6
.2 NOTE: Symbols as in Fig. 4-43
oL---~---~---L---~---~----~0 Fig. 4-45. Specific weight of
turboprop engines (data from Ref. 4-27)
50 100 150 200 250 300 350
I . HP
SPECIFIC SHAFT HORSEPOWER Fj0 W- KG/SEC
M0 .4 at 19,680 ft (6000 m), resulting in Fig. 4-42, which shows the following:
a. The specific fuel consumption continues to decrease until a pressure ratio of ap-proximately 8 is reached; beyond this little gain can be expected~
b. The influence of the TET on fuel con-sumption is not very great. The value Cp=
. 45 lb/hp/h (.20 kg/hp/h) may be regarded as a practical lower limit for the
condi-*Greater pressure ratios may still be chosen for engines which should have a low fuel consumption at lower rpm than the de-sign condition.
tions considered.
c. For each OPR the specific power rises considerably with increasing TET. Moreover, at low OPR, with a given TET, the specific power increases with increasing OPR. How-ever, for each TET an optimum OPR is found, where the specific power will be maximum.
Fig. 4-43 shows the combinations of OPR and TET which have been used in actual engines • d. Contrary to the situation found for the jet engine, increasing TET results in ever decreasing specific fuel consumption, the reason being that the propeller efficiency is independent of the TET, whereas the thermal efficiency improves with increasing
TET.
Fig. 4-43 shows that for a number of en-gines the values adopted for the OPR do not depart to any great extent from the calculated optimum. This leads us to think that further development of turboprop en-gines will not prove spectacular as far as specific fuel consumption and specific power are concerned. There are, however, still possibilities of achieving a low fuel consumption for engines which have to run at low power during very long flights, e.g.
regenerative engines, for which the reader should consult Ref. 4-6.
The effect of the OPR on specific fuel ·con-sumption is shown statistically in Fig.
4-44.
4.5.2. Weight and drag
The weight of the turboprop engine is largely decided by the mass of air flowing through it and amounts to about 45 lb per lb/s of airflow. Therefore:
we 45
- = - - -
(lb/hp or kg/hp)Pto (P/W)to (4-40)
From this expression, which is compared with actual examples in Fig. 4-45, it is concluded that a high specific power is important for low weight. To this we may add that the frontal area of the engine is also proportional to W, so for the nacelle drag per hp it is possible to derive a similar relationship, as in (4-40).
4.5.3. Turboprop engine configurations Older types of turboprop engines are fitted with a single centrifugal compressor. (See Fig. 4-46). SinceOPR values of about 4 have been achieved, these engines show a specific
fuel consumption of at least .65 to .75 lb/hp/h (.30 to .35 kg/hp/h). For this reason the air is often pre-compressed by means of an axial compressor and with this combination it will be possible to reach pressure ratios of the order of 10. These
pressure ratios may also be obtained with a single axial compressor or with two ctrifugal compressors in series. Some en-gines are fitted with two axial compressors on separate shafts and this makes a high pre1sure ratio possible (e.g. Rolls-Royce Tyne, OPR • 13.5, see Fig. 4-46b).
Many turboprop engines are fitted with a free turbine. This is a low pressure tur-bine which is not coupled mechanically to the gas generator, but drives the propeller through a separate shaft·. With these en-gines it is often possible for the pilot
·to adjust the power in certain phases of the flight by selecting the propeller pitch
(8-control). This demands a special device which governs the quantity of fuel supplied to the eng~ne and controls the power output in such a way that the propeller rpm re-mains constant (cf. Section 6.3.3.).
The following features are offered by the free turbine:
a. When output has to be rapidly increased in a flight phase with low power, it will only be necessary to increase the rpm of the gas generator. The high-inertia pro-peller already revolves at the required rpm, thus making it possible to obtain a quick response.
b. In flight conditions which require con-trol of the aircraft relative to a given glide path the use of 8-control will give improved speed stability, particularly in the low speed regime.
c. Due to the aerodynamic coupling with the gas generator, the free turbine runs close to the optimum rpm under different working conditions,.resulting in a high efficiency.
d. In the case of engine failure, the free-ly revolving propeller will onfree-ly have to drive the free turbine and propeller drag immediately after the failure will be small.
There are various possible engine cvnfigu-rations which can be adapted to feature a free turbine, without resorting to the co-axial layout. Examples are the Allison 250
(Fig. 4-46c) and the Pratt and WhitneyPT6A, 137
a. Rolls-Rovce DART-Mk . SIO
b. Rolls-Royce TYNE 12
c. Allison Model 250
138
Fig. 4-46 . Examples of turboprop engine configurat-ions (data can be found in Table 6-2)
d. Pratt and Whitney PT-6A
e. Turbom~ca Astazou XIV
Fig. 4-46. Examples of turboprop engine configurations. (data can be found in Table 6-2)
(Fig. 4-46d), which work on the "reverse flow" principle with reversal of the flow within the engine. The gases are ejected either in front or in the centre of the engine at the sides. The advantage lies in
the very compact layout, demanding little space to accommodate the engine. Thi~ may be compared with the Turbom~ca Astazou
(Fig. 4-46e) with the jet pipe at the rear of the engine. However, with the jet pipea at the sides it is not possible to take full advantage of the thrust of the engine gases, unless one or more curved jet pipea are used . This will lead to an increase in drag and a loss of thrust.
139