GB2238080A - Propulsion system for an aerospace vehicle - Google Patents

Abstract

A hybrid propulsion system for a transatmospheric or hypersonic cruise vehicle comprises a pre-cooled air-breathing jet engine 10 and a rocket motor 12. Previously ice formed on the inlet cooling heat exchanger 40 at altitudes where there is significant humidity sufficiently to deleteriously affect jet thrust. The rocket motor is therefore operated temporarily to compensate for the loss of jet thrust when icing is present and during defrost. The rocket motor is mounted co-axially with the jet pipe which exhausts through an annular orifice 36 into a common propelling nozzle 16. The arrangement avoids transverse thrust vectors as the relative engine thrusts vary which would cause the vehicle to pitch or yaw. The air inlet heat exchange is intermittently by-passed via line 50 to allow deicing. <IMAGE>

Description

IMPROVED PROPULSION SYSTEM FOR AN AEROSPACE VEHICLE The invention concerns improvements to propulsion systems of the kmd'omprising a rocket motor and a pre-cooled turbo-jet.

Propulsion systems of this kind are proposed for operation throughout the flight envelope of hypersonic cruise and trans-atmospheric aerospace vehicles, from ground level take-off to orbital insertion. Although rocket engines have speed limitations and are free to operate at all altitudes, in general, they are possessed of a very high specific fuel consumption about ten times higher, comparatively, than a gas turbine engine.

Therefore, mixed propulsion systems have been proposed to make better use of an aircraft fuel load by running a gas turbine engine at relatively low speeds and low altitudes, and reserving the rocket engine for use at higher speeds and altitudes where the jet performance tails-off.

A straight forward air breathing gas turbine is limited to about Mach 3.5 by the very high equivalent temperature of the airstream in the air inlet. Pre-cooling the airstream by means of a heat exchanger in the air inlet increases the max flying speed of the engine to about Mach 6. At higher speeds than this the rocket motor takes over and is used exclusively. The common fuel, liquid hydrogen, used in both engines has very large thermal capacity and by partially expanding the engine supply fuel in the heat exchanger it can be used to absorb much of the kinetic energy of the intake air.

Icing of the heat exchanger is inherent in this kind of system as the humidity of the air rises. A solution to this problem is to allow the cooling fuel flow to by-pass the heat exchanger matrix temporarily and thereby enable the airstream to de-ice the matrix. This has the drawback that engine thrust varies in accordance not only with the degree of icing but also as a result of the interruption of the cooling cycle. The present invention proposes to overcome this drawback.

According to the present invention there is provided a propulsion system of the kind comprising a rocket motor and a pre-cooled turbo-jet engine in which air entering the jet is cooled by passage through a heat exchanger, and the rocket motor and the jet engine have control systems interconnected such that the rocket motor is run in the regime of the air-breathing engine at a thrust rating variable to compensate for thrust losses in the air-breathing engine.

The rocket motor may be operated selectively at variable thrust ratings to compensate for a reduction in jet thrust arising from variations in the air intake pre-cooler temperature drop due to ice accretion in the heat exchanger matrix, to produce extra thrust in critical flight conditions, and to effect a progressive change-over to all-rocket propulsion towards the limit of the air-breathing engine regime.

Preferably the cooling fluid for the heat exchanger comprises a vaporising liquid fuel common to both engines and to initiate de-icing of the heat exchanger a by-pass path for the fuel.

The invention will now be described in more detail, by way of example only, with reference to the single figure of the accompanying drawing which shows diagrammatically a mixed air breathing and rocket propulsion system.

The illustrated propulsion system comprises an air-breathing single-spool axial flow turbo-jet engine 10 and a rocket motor 12. In so far as both engines are symmetrical about a centre axis, they are mounted co-axially, with the rocket motor to the rear of the jet engine. The air inlet for the turbo-jet is generally indicated at 14, on the left side of the drawing, and both engines have a common exhaust nozzle, generally indicated at 16, towards the right in the drawing.

The jet engine 10 illustrated in the accompanying drawings, purely for the purposes of exemplary description is of the axial flow, single spool type having typically an axial flow compressor 18 providing a compression ratio of 20:1 and supplying air to an annular combustor 20 which, in turn, feeds a two stage turbine 22. It is to be understood that the invention is not restricted in respect of the specific type of air-breathing engine configuration or engine cycle. The turbine exhausts into a jet pipe 24 containing a re-heat or afterburning system 26. The pipe 24 terminates in a convergent-divergent variable area propelling nozzle 28. The nozzle 28 is capable of being contracted to a position in which the jet pipe 24 is completely shut off in which position the divergent nozzle effectively forms a continuation of the walls of the rocket propelling nozzle.

Disposed within the jet pipe 24 is the rocket motor 12 which is mounted co-axially with the jet engine 24 so that the two engines deliver thrust on the same axis with no offset. The motor 12 has a combustion chamber 30 from which, in operation, expanding rocket gas passes into a second convergent-divergent propelling nozzle 32. This second nozzle is supported concentrically with the first nozzle and has a fixed area exit orifice 34 lying in substantially the same plane as the throat 36 of the first nozzle 28.

The turbo-jet engine 10 draws in air through an aerodynamic duct 38 containing a heat exchanger 40 supplied with high pressure liquid hydrogen which pre-cools the air entering the compressor 18.

Typically the temperature range of air entering the duct 38 is 250-12500K and when entering the compressor 18, after passing through the heat exchanger 40, the temperature has been cooled to temperature with the range 100-2800K 1500K is a typical jet entry temperature but still lower temperatures, say down to 100or may be found advantageous in some circumstances.

Liquid hydrogen is taken from the main fuel tanks (not shown) by a turbine pump 42 and its pressure raised substantially, typically to several hundred bar, by compressors 44 and 46. The high pressure outlet of compressor 46 is connected to the heat exchanger 40 through a cut-off valve 48. A by-pass circuit around the heat exchanger is formed by a parallel connection 50 containing a second cut-off valve 52. Hydrogen from the outlet of heat exchanger 40 or from the by-pass circuit 50 is delivered to a further heat exchanger 54 built into the rocket combustion chamber 30. The connection 50 is also branched to supply cooling hydrogen to a further heat exchanger 56 in the propelling nozzle 32. The hot hydrogen is fed back to power turbine 58 driving the compressors 44, 46 and the pump 42.

The output of turbine 58 is divided between two further circuits. The first of these supplies gaseous hydrogen fuel to the turbo-jet combustor manifold 20 through a first control valve 60 and through a second control valve 62 to the reheat system 26. The second of the turbine outputs drives a further turbine 64 coupled to a liquid oxygen compressor 66. Liquid oxygen is drawn from an oxident supply tank (not shown) through a first compressor stage 68 driven by a turbine 70 which supplies the input to compressor 66 and is driven by one of two outputs from the compressor. The second output from compressor 66 supplies oxident to rocket motor. The hydrogen exiting from turbine 64 is passed back to the combustion chamber 30 as a fuel supply, it may also pass directly via by-pass connections including one-way valves 66, 68 around turbines 58 and 64 respectively.

In operation of the illustrated arrangement the turbo-jet engine 10 is used preferentially to power the vehicle at low altitudes and take-off. The reheat system 26 may be used as required to provide additional thrust, for example, at take off and for acceleration to supersonic speed. In normal flight the jet engine is operated at constant rpm. The rocket motor 12 is preferably reserved for use at high altitudes where there is insufficient or no atmosphere to supply the air-breathing turbo-jet.

However, in accordance with the invention the rocket motor is run in an idle condition at moderate altitudes below about 30,000 ft where icing conditions may be encountered, in readiness for immediate use as a supplementary source of thrust. There is a range of altitudes up to about 30,000 ft above which there is sufficient humidity in the atmosphere to cause heavy ice accretions to form in the heat exchanger 40 as a result of which the turbo-jet inlet becomes restricted. The jet engine therefore loses thrust. The heat exchanger must be de-iced to restore lost thrust and to accomplish this the hydrogen coolant is made to by-pass the heat exchanger 40 by switching over valves 48 and 52. The kinetic energy of the air flowing through the jet inlet soon clears the ice and engine thrust is restored. This cycle of events will probably occur repeatedly.

During the periods of lost or reduced power additional fuel is burned in the rocket motor to produce extra thrust to compensate for losses experienced in the jet engine. The engine control is arranged to vary the levels of fuel supplied to the jet and rocket engines accordingly in order to maintain a substantially smooth power curve verses time. Changes in the fuel supply levels to both engines are preferably blended to avoid abrupt changes and the control systems are arranged to provide a progressive change-over from air-breathing to rocket mode towards the upper limit of the jet engine. Means operative in the dual engine mode is provided for sensing the amount of ice deposited in the heat exchanger and initiating a de-icing operation. Simultaneously the fuel supply to the rocket motor is increased to produce extra thrust to compensate for the reduced efficiency of the jet engine.Also, the control systems for the two engines may be arranged to permit use of the rocket motor to produce additional thrust during critical flight conditions, for example, at transonic speeds rather in the manner that a reheat system might be used in a conventional aircraft. Use of the rocket motor in the jet engine regime is, therefore, not limited to situations where the jet thrust is degraded.

Disposition of the rocket motor co-axially with the jet obviates the need to retrim the vehicle due to changes in the balance of the thrust vectors of the two engines when the rocket thrust is increased to supplement jet thrust or to compensate for lost jet thrust. By mounting the two engines co-axially thrust components are avoided which would otherwise cause a tendency for the vehicle to pitch, if the engines were displaced vertically, or to yaw, if the engines were displaced laterally. Also, the mounting arrangement permits the use of a common propelling nozzle for the two engines. The rocket motor combustion chamber nozzle is surrounded by an annular jet pipe duct. In the pure rocket mode the walls of the propelling nozzle are contracted adjacent the end of the jet pipe to provide a relatively smooth and continuous wall to the propelling nozzle.

Claims (11)

1. A propulsion system of the kind comprising a rocket motor and a pre-cooled turbo-jet engine in which air entering the jet is cooled by passage through a heat exchanger, and the rocket motor and jet engine have control systems interconnected such that the rocket motor is run in the regime of the air-breathing engine at a variable thrust rating to supplement thrust produced by the air-breathing engine.

2. A propulsion system as claimed in claim 1 wherein the air-breathing engine and rocket motor are disposed to produce thrust substantially on a common axis.

3. A propulsion system according to claim 1 wherein the jet and rocket engines share a common propulsion exhaust nozzle.

4. A propulsion system according to claim 3 wherein the jet exhausts in the nozzle through an annular duct placed concentrically with the rocket exhaust.

5. A propulsion system according to claim 3 or 4 wherein the jet exhaust nozzle is arranged to be closed-off in a rocket only mode.

6. A propulsion system according to any preceding claim wherein the heat exchanger in the jet air intake is cooled by partially expanded fuel.

7. A propulsion system according to claim 6 wherein the heat exchanger is cooled by hydrogen.

8. A propulsion system according to claim 6 or 7 further comprising means responsive to a predetermined level of ice accretion in the heat exchanger to initiate a de-icing operation and the engine control systems operate to increase rocket thrust to compensate for temporary reduction of thrust produced by the air-breathing engine.

9. A propulsion system according to claim 8 wherein the heat-exchanger is provided with a by-pass circuit to divert coolant for a de-icing operation.

10. A propulsion system according to any preceding claim further comprising control means arranged to effect a progressive change-over from the air-breathing engine mode to the rocket mode.

11. A propulsion system substantially as described with reference to the accompanying drawing.