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0
g
(P P ) A W VJ
= − 0 ⋅ + −
0
32
= (6 − 0) ⋅ 332 + 153 X 1,917 −
Hypotenuse
Base
system are reheated to provide additional thrust. The
effect of afterburning is to increase the volume of the
exhaust gases, thus producing a higher exit velocity
at the propelling nozzle.
24. Assuming that an afterburner jet pipe and
propelling nozzle are fitted to the engine used in the
previous calculations, and the new conditions at the
propelling nozzle are as follows-
OUTLET Area (A) = 455 sq.in.
Pressure (P) = 5 lb. per sq.in.
(gauge)
Velocity (vJ) = 2,404 ft. per sec.
Mass flow (W) = 157 lb. per sec.
The thrust
= 14,069 - 16,745
= 2,676 lb. acting in a rearward direction.
Therefore, compared with the previous calculation in
para. 17, it will be seen that the negative thrust is
reduced from -5,587 lb. to -2,676 lb.; the overall
positive thrust is thus increased by 2,911 lb; which is
equivalent to a thrust increase of more than 25 per
cent.
25. To arrive at the total thrust of the engine with
afterburning the calculations in para. 20 should use
the above figures.
Thrust distribution
213
16,745
g
(A x P) W VJ = + −
16,745
32
= (455 x 5) + 157 x 2,404 −
Rolls-Royce RB168 MK807
Blackburn Nimbus
The Nimbus was developed from the A129
turbo-shaft which, in its turn, was a modified
Turbomeca Artouste built under licence. The
Nimbus developed 968 hp, but for helicopter
use was flat-rated at 710 hp. The engine was
used in Westland Wasp and Scout helicopters
and four 700 hp units were used to power the
experimental 5RN-2 hovercraft.
21: Performance
Contents Page
Introduction 215
Engine thrust on the test
bench 217
Comparison between thrust and
horse-power
Engine thrust in flight 218
Effect of forward speed
Effect of afterburning on engine
thrust
Effect of altitude
Effect of temperature
Propulsive efficiency 223
Fuel consumption and
power-to-weight relationship 225
INTRODUCTION
1. The performance requirements of an engine are
obviously dictated to a large extent by the type of
operation for which the engine is designed. The
power of the turbo-jet engine is measured in thrust,
produced at the propelling nozzle or nozzles, and
that of the turbo-propeller engine is measured in
shaft horse-power (s.h.p.) produced at the propeller
shaft. However, both types are in the main assessed
on the amount of thrust or s.h.p. they develop for a
given weight, fuel consumption and frontal area.
2. Since the thrust or s.h.p. developed is dependent
on the mass of air entering the engine and the acceleration
imparted to it during the engine cycle, it is
obviously influenced, as subsequently described, by
such variables as the forward speed of the aircraft,
altitude and climatic conditions, These variables
influence the efficiency of the air intake, the
compressor, the turbine and the jet pipe; consequently,
the gas energy available for the production
of thrust or s.h.p. also varies.
3. In the interest of fuel economy and aircraft range,
the ratio of fuel consumption to thrust or s.h.p. should
be as low as possible. This ratio, known as the
specific fuel consumption (s.f.c.), is expressed in
pounds of fuel per hour per pound of net thrust or
s.h.p. and is determined by the thermal and
propulsive efficiency of the engine. In recent years
considerable progress has been made in reducing
s.f.c. and weight. These factors are further explained
in para. 46.
215
4. Whereas the thermal efficiency is often referred
to as the internal efficiency of the engine, the
propulsive efficiency is referred to as the external
efficiency. This latter efficiency, described in para. 37,
explains why the pure jet engine is less efficient than
the turbo-propeller engine at lower aircraft speeds
leading to development of the by-pass principle and,
more recently, the propfan designs.
5. The thermal and the propulsive efficiency also
influence, to a large extent, the size of the
compressor and turbine, thus determining the weight
and diameter of the engine for a given output.
6. These and other factors are presented in curves
and graphs, calculated from the basic gas laws (Part
2), and are proved in practice by bench and flight
testing, or by simulating flight conditions in a high
altitude test cell. To make these calculations, specific
symbols are used to denote the pressures and temperatures
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