GB2624129A

GB2624129A - Rocket engine

Info

Publication number: GB2624129A
Application number: GB2401822.8A
Authority: GB
Inventors: J P Wirth Nicholas
Original assignee: Wirth Research Ltd
Current assignee: Wirth Research Ltd
Priority date: 2024-02-09
Filing date: 2024-02-09
Publication date: 2024-05-08
Also published as: GB202401822D0

Abstract

A rocket engine 2 comprising a combustion chamber 4 and a first pump having a first rotor assembly for pumping a first propellant to the combustion chamber 4. Wherein a portion of the rotor assembly is annular and is arranged to encircle and rotate relative to the combustion chamber. The rocket may comprise a second pump having a second rotor assembly 60, 62, 64 for pumping a second propellant to the combustion chamber. The first propellant may be fuel, for example liquid hydrogen, and the second propellant may be oxidiser, for example liquid oxygen. A portion of the second rotor may be annular and arranged to encircle and rotate relative to the combustion chamber. The first and second pump may be arranged in opposite directions.

Description

ROCKET ENGINE

The present application relates to a rocket engine and particularly, but not exclusively, to a rocket engine with a concentric turbopump.

BACKGROUND

A major hurdle for increasing the frequency of space travel is the cost. The cost of space travel can be improved in two ways: increasing the payload carrying capability of a rocket of given size, or reducing the mass of a rocket with a given payload.

Tsiolkovsky's Rocket Equation is a simple method of determining the ability of a rocket to reach orbit with a given payload.

AV is the change in velocity across the rocket burn and is the fundamental measure in the performance of any rocket -the higher AV the better. /s" is the specific impulse of the rocket engine and propellant choice which defines the efficiency; again, a higher value is better. Mi"itica is the initial mass of the fully fuelled rocket before take-off, whereas Mii""/ is the final mass of the rocket when all propellant has been burnt and thus corresponds to the mass of the rocket structure plus the mass of the payload.

Therefore, reducing Miimil provides a greater AV, and thus a lighter structure for a given payload needs to burn less fuel to reach orbit.

It is therefore desirable to develop a rocket engine architecture that reduces weight in order to reduce the fuel costs and also the cost and time to manufacture the engine.

SUMMARY

In accordance with an aspect of the invention, there is provided a rocket engine comprising: a combustion chamber; a first pump comprising a rotor assembly for pumping a first propellant to the combustion chamber; wherein at least a portion of the rotor assembly is annular and is arranged so as to encircle and rotate relative to the combustion chamber.

The rocket engine may further comprise a second pump comprising a rotor assembly for pumping a second propellant to the combustion chamber.

At least a portion of the rotor assembly of the second pump may be annular and arranged so as to encircle and rotate relative to the combustion chamber.

The rotor assembly of the first pump may overlap in an axial direction with the rotor assembly of the second pump.

The first pump and the second pump may be arranged in opposite directions.

The rotor assembly of the first and/or second pump may comprise an impeller. The impeller may be annular and arranged so as to encircle and rotate relative to the combustion chamber.

The rotor assembly of the first and/or second pump may further comprise a turbine configured to drive the first and/or second pump.

The turbine may be annular and arranged so as to encircle and rotate relative to the combustion chamber.

The first pump is fluidically connected to the combustion chamber via a first propellant channel.

The first pump may be fluidically connected to a chamber cooling channel which surrounds the combustion chamber and extends axially along its length.

The chamber cooling channel may be fluidically connected to a turbine drive channel downstream of the combustion chamber. The turbine of the first and/or second pump may be disposed within the turbine drive channel.

The turbine drive channel may comprise a bypass duct and a bypass control valve which allows the first propellant to bypass the turbine of one or both of the first and second pumps.

The turbine drive channel may be fluidically connected to a turbine exhaust channel downstream of the turbine of the first and/or second pump. The turbine exhaust channel may extend alongside the first propellant channel so as to provide heat transfer therebetween.

A portion of the turbine exhaust channel may be radially outward of the first propellant channel and a portion of the turbine exhaust channel may be radially inward of the first propellant channel. The turbine exhaust channel and the first propellant channel may each comprise a plurality of channel portions which are interleaved to allow the turbine exhaust channel to cross over the first propellant channel.

The turbine exhaust channel may terminate within a nozzle of the engine.

The turbine exhaust channel may terminate downstream of a throat of the nozzle.

The annular portion of the rotor assembly may be concentric with the combustion chamber.

The first propellant may be fuel and the second propellant may be oxidizer.

The first propellant may be liquid hydrogen and the second propellant may be liquid oxygen.

BRIEF DESCRIPTION OF DRAWINGS

For a better understanding of the invention, and to show more clearly how it may be carried into effect, reference will now be made, by way of example, to the accompanying drawings, in which: Figure 1 is a perspective view of a rocket engine according to an embodiment of the 35 invention; Figure 2 is a cross-sectional view of the rocket engine; and Figures 3 and 4 are cross-sectional view showing an area where a turbine exhaust channel crosses a first propellant channel.

DETAILED DESCRIPTION

Figures 1 and 2 show a rocket engine 2 according to an embodiment of the invention.

The rocket engine 2 generally comprises a combustion chamber 4 which is fluidically connected to a nozzle 6. The combustion chamber 4 is formed by a cylindrical wall.

The nozzle 6 comprises a convergent section 8 which converges from the combustion chamber 4 to a throat 10 and a divergent section 12 which diverges away from the throat to a nozzle exit 14 at the distal end of the nozzle 6. The nozzle 6 may be a Rao nozzle (or Thrust Optimised Parabolic nozzle).

The rocket engine 2 further comprises a first tank 16 containing a first propellant and a second tank 18 containing a second propellant. In this example, the first propellant is a fuel, particularly liquid hydrogen, and the second propellant is an oxidizer, particularly liquid oxygen.

The first tank 16 is connected to a first duct 20. As shown in Figure 1, the first duct 20 comprises an axial portion 22 which extends along a longitudinal (rotational) axis of the rocket engine 2 and a circumferential portion 24 which extends circumferentially about the combustion chamber 4.

The rocket engine 2 further comprises a first turbopump. The first turbopump generally comprises a first centrifugal pump 26 and a first turbine 28 which are mounted on a common first shaft 30 to form a first rotor assembly. Specifically, the first centrifugal pump 26 comprises an inducer 32 and three impellers 34a, 34b, 34c (forming three stages) which are disposed in series downstream of the inducer 32. The impellers 34a, 34b, 34c are located within an outer casing 39. A diffuser 36a, 36b, 36b is provided downstream of each of the respective impellers 34a, 34b, 34c. The circumferential portion 24 of the first duct 20 forms an inlet volute which is in fluid communication with an inlet of the first centrifugal pump 26 at the inducer 32.

The first centrifugal pump 26 is configured to receive and pressurise the first propellant (e.g., liquid hydrogen) from the first tank 16, as will be described in further detail below.

An outlet of the first centrifugal pump 26 is in fluid communication with a first propellant channel 38 via a first control valve 40 (Main Fuel Valve -MR/) and with a chamber cooling channel 42 via a second control valve 44 (bleed control valve).

The chamber cooling channel 42 is provided around the combustion chamber 4. In this example, the chamber cooling channel 42 is formed between the wall of the combustion chamber 4 and a further cylindrical wall spaced radially outwards from the wall of the combustion chamber 4. The chamber cooling channel 42 extends longitudinally along the length of the combustion chamber 4. In this example, the chamber cooling channel 42 also extends partially along the length of the nozzle 6. In other examples, the chamber cooling channel 42 may extend further or entirely along the length of the nozzle 6. The chamber cooling channel 42 extends around the circumference of the combustion chamber 4 to form an annular jacket. A plurality of longitudinally extending ribs 43 are provided in the chamber cooling channel 42 which divide the chamber cooling channel 42 into a plurality of annular sectors over at least part of the longitudinal length of the chamber cooling channel 42.

The first propellant channel 38 extends along the length of the combustion chamber 4, radially outward of the chamber cooling channel 42. In this example, the first propellant channel 38 also extends partially along the length of the nozzle 6. The first propellant channel 38 extends around the circumference of the combustion chamber 4. Like the chamber cooling channel 42, a plurality of longitudinally extending ribs 46 are provided in the first propellant channel 38 which divide the first propellant channel 38 into a plurality of annular sectors over at least part of the longitudinal length of the first propellant channel 38.

The first propellant channel 38 is fluidically connected to an injector assembly 48 provided at the top of the combustion chamber 4 via a baffle 50.

The pressurised first propellant from the first centrifugal pump 26 is fed along the first propellant channel 38 in an upward direction (i.e. away from the nozzle exit 14) to the injector assembly 48. The flow of first propellant can be controlled using the first control valve 40.

A portion of the pressurised first propellant can also be diverted to the chamber cooling channel 42 so as to provide cooling of the wall of the combustion chamber 4.

The chamber cooling channel 42 is fluidically connected to a turbine drive channel 52. In some examples, the chamber cooling channel 42 and turbine drive channel 52 may be continuous (i.e., integrally formed). The turbine drive channel 52 returns along the length of the combustion chamber 4, radially outward of the chamber cooling channel 42. The first turbine 28 of the first turbopump is disposed within the turbine drive channel 52. Specifically, a plurality of turbine blades 54 extend radially inward from the first shaft 30 and are spaced around the circumference of the turbine drive channel 52.

The pressurised, heated and vapourised first propellant passes along the turbine drive channel 52 after exiting the chamber cooling channel 42 and drives the first turbine 28. This rotates the first shaft 30 which in turn rotates the first centrifugal pump 26 in order to pressurise the first propellant.

The rocket engine 2 further comprises a second turbopump. The second turbopump generally comprises a second centrifugal pump 60 and a second turbine 62 which are mounted on a common second shaft 64 to form a second rotor assembly. Specifically, the second centrifugal pump 60 comprises an inducer 66 and an impeller 68 downstream of the inducer 66. The impeller 68 is located within an outer casing 70.

The outer casing 70 is formed as a bell housing.

The second tank 18 is connected to the second centrifugal pump 60 via a second duct 72. As shown in Figure 1, the second duct 72 comprises an axial portion 74 which extends along the longitudinal axis of the rocket engine 2 and a circumferential portion 76 which extends circumferentially about the combustion chamber 4. The circumferential portion 76 of the second duct 72 forms an inlet volute which is in fluid communication with an inlet of the second centrifugal pump 60 at the inducer 66.

The second turbine 62 of the second turbopump is also disposed within the turbine drive channel 52. Specifically, a plurality of turbine blades 78 extend radially inward from the second shaft 64 and are spaced around the circumference of the turbine drive channel 52. The second turbine 62 is located upstream of the first turbine 28.

An inlet stator 56 is provided within the turbine drive channel 52 upstream of the second turbine 62. The inlet stator 56 comprises a plurality of stator blades 58 spaced around the circumference of the turbine drive channel 52.

Further, an intermediate stator 80 is provided within the turbine drive channel 52 between the second turbine 62 and the first turbine 28. The intermediate stator 80 comprises a plurality of stator blades 82 spaced around the circumference of the turbine drive channel 52.

The second centrifugal pump 60 is configured to receive and pressurise the second propellant (e.g., liquid oxygen) from the second tank 18.

As described previously, the pressurised, heated and vapourised first propellant passes along the turbine drive channel 52 after exiting the chamber cooling channel 42. In addition to driving the first turbine 28 and in turn the first centrifugal pump 26, the pressurised, heated and vapourised first propellant also drives the second turbine 62.

This rotates the second shaft 64 which in turn rotates the second centrifugal pump 60 and thereby pressurises the second propellant.

The pressurised second propellant is directed from the outlet of the second centrifugal pump 60 through a second propellant channel 84 formed within the outer casing 70 and to the injector assembly 48 where it is introduced via an injector array 86. A third control valve 92 (Main Oxygen Valve -MOV) is provided in the second propellant channel 84 which allows the flow of second propellant to be controlled.

The first and second shafts 30, 64 are each cylindrical (i.e., tubular) and are supported by bearings. Specifically, the first shaft 30 is supported by first and second sets of bearings 88a, 88b disposed about its external surface and spaced longitudinal from one another. The second set of bearings 88b are located between the first shaft 30 and the second shaft 64. Similarly, the second shaft 64 is supported by first and second sets of bearings 90a, 90b disposed about its external surface and spaced longitudinal from one another. The first and second shafts 30, 64 are therefore independently rotatable relative to one another and relative to the combustion chamber 4.

A bypass control valve 94 is provided along the turbine drive channel 52. The bypass control valve 94 is located adjacent to the intermediate stator 80 and between the first turbine 28 and the second turbine 62. The bypass control valve 94 may be selectively opened to allow the pressurised, heated and vapourised first propellant to bypass the first turbine 28 via a bypass duct 96. The bypass control valve 94 can therefore be used to vary the speed of rotation of the first turbine 28 independent of the second turbine 62 (and thus the speed of the first centrifugal pump 26 independent of the second centrifugal pump 60). The bypass control valve 94 may be formed as a variable blow-off ring. For example, the variable blow-off ring may comprise a plurality of apertures spaced around its circumference and the turbine drive channel 52 may comprise a plurality of corresponding apertures. The rotational position of the variable blow-off ring can therefore be adjusted so that the apertures are aligned with one another (in a fully open position) or are offset (or partially offset) so that the apertures in the turbine drive channel 52 are blocked by portions of the variable blow-off ring between adjacent apertures.

The turbine drive channel 52 is connected to a turbine exhaust channel 98 downstream of the first turbine 28. In some examples, the turbine drive channel 52 and turbine exhaust channel 98 may be continuous (i.e., integrally formed). Like the first propellant channel 38 and the chamber cooling channel 42, a plurality of longitudinally extending ribs 102 are provided in the turbine exhaust channel 98 which divide the turbine exhaust channel 98 into a plurality of annular sectors over at least part of the longitudinal length of the first propellant channel 38.

As shown in Figures 3 and 4, the turbine exhaust channel 98 crosses the first propellant channel 38. Specifically, the turbine exhaust channel 98 comprises a plurality of discrete channel portions and the first propellant channel 38 comprises a plurality of discrete channel portions. The channel portions of the turbine exhaust channel 98 are interleaved with the channel portions of the first propellant channel 38. In an upward direction, the first propellant channel 38 therefore deviates radially inwardly and the turbine exhaust channel 98 deviates radially outwardly. The turbine exhaust channel 98 is therefore disposed between the chamber cooling channel 42 and the first propellant channel 38 along a lower portion. The channel portions of the turbine exhaust channel 98 and the channel portions of the first propellant channel 38 may be defined by the ribs 102, 46 respectively.

As shown, the first propellant channel 38 also crosses the turbine drive channel 52 upstream of the injector assembly 48. A similar arrangement of interleaved channel portions as described above is also used here to allow this.

The first propellant flowing through the turbine exhaust channel 98 has been heated relative to the first propellant flowing through the first propellant channel 38 and thus acts to warm the first propellant before it enters the injector assembly 48. In particular, the turbine exhaust channel 98 and the first propellant channel 38 may share a common wall along the lower portion in order to allow heat transfer. This improves performance and extracts useful energy from the turbine exhaust flow. Likewise, the chamber cooling channel 42 and the turbine exhaust channel 98 may share a common wall to allow heat transfer. As can be seen, the flow along the turbine exhaust channel 98 is in the opposite direction (downwards) compared to the flow in the chamber cooling channel 42 and the first propellant channel 38 (upwards).

The turbine exhaust channel 98 terminates at an exit port 100 located within the nozzle 6. The exit port 100 is located downstream of the throat 10 such that the exhaust is at a very low pressure point. As shown, the turbine exhaust channel 98 crosses the chamber cooling channel 42 upstream of the exit port 100. A similar arrangement of interleaved channel portions as described above is also used here to allow this.

As described previously, the first propellant and the second propellant are supplied to the injector assembly 48. The second propellant is injected longitudinally via the injector array 86, and the first propellant flows through this region after passing through the baffle 50, and is then also injected longitudinally through the injector plate 87. The injector assembly 48 comprises a spark ignitor 104 which ignites the first and second propellants in ignition channel 105, with this leading to main propellant ignition in the combustion chamber 4 below, with the exhaust gases being expelled from the nozzle 6 to create thrust.

The rocket engine 2 uses the open expander cycle which utilises a "bootstrapping" start sequence. Specifically, to start the rocket engine 2, the first, second and third control valves 40, 44 and 92 are opened. The first propellant is held in the first tank 16 under pressure and so causes the first centrifugal pump 26 to rotate once released. The spark ignitor 104 is activated so as to commence combustion. This heat starts to add energy to the first propellant flowing through the chamber cooling channel 42 and thus starts to rotate the first and second turbines 28, 62 which increases the flow through the first and second centrifugal pumps 26, 60.

Once running pressure has been achieved, the power level and mixture ratio of first and second propellants may be adjusted by varying the position of the control valves.

As described, the rocket engine 2 utilises rotor assemblies within the first and/or second turbopumps which have at least a portion which encircles the combustion chamber and rotates about the combustion chamber. In particular, at least a portion of at least one of the shaft, centrifugal pump and turbine of the rotor assembly is annular and encircles the combustion chamber.

As described above, the first propellant channel 38 and the turbine exhaust channel 98 are arranged to provide heat exchange. The first propellant channel 38 and the turbine exhaust channel 98 are located radially inward of the rotor assemblies and thus are between the combustion chamber 4 and the rotor assemblies. In other examples, the first propellant channel 38 and the turbine exhaust channel 98 may be routed so as to be radially outward of the rotor assemblies.

In the example shown, the rotational axis of each rotor assembly is concentric with a central axis of the combustion chamber. However, in other examples the rotational axis may be offset from the axis of the combustion chamber; for a plurality of separate combustion chambers may be utilised. Further, in other examples, the or each combustion chamber may not be cylindrical. For non-cylindrical combustion chambers, the rotational axis of each rotor assembly may be centred on the centroid or centre of mass of the cross-sectional shape or may be offset from any such point.

The impellers may be of axial, centrifugal or mix-flow design or any combination. The inducers are intended to prevent cavitation, but may not be required in some instances. Boost pumps may also be used prior to the propellant reaching the main turbopump assembly.

In the example shown, the first and second centrifugal pumps are arranged in opposite directions. In particular, the first centrifugal pump 26 is arranged such that its inlet is located above its outlet, whereas the second centrifugal pump 60 is arranged such that its inlet is below its outlet. Moreover, the inlets of the first and second centrifugal pumps are adjacent to one another (towards the centre of the rocket engine 2 in an axial direction). This arrangement provides a particularly efficient use of space. However, in other examples, the first and second centrifugal pumps may be arranged in the same direction.

Further, in the example shown, the first and second turbopumps overlap one another in order to reduce space. However, in other examples, the first and second turbopumps may be arranged entirely in series, with no overlap of their respective rotor assemblies in an axial direction.

Although the rocket engine 2 has been described as using a pair of turbopumps, it will be appreciated that any number of turbopumps may be used. For example, in other examples, only a single turbopump may be used (e.g., with a single propellant). Further, in other examples, the rocket engine 2 may use three or more propellants (such as two or more fuels plus an oxidizer). Accordingly, where necessary, separate turbopumps may be provided for each propellant. Alternatively, multiple propellants may be pressurised by a single pump.

Furthermore, a similar integration may also be beneficial with other types of pump. For example, the engine may use a centrifugal pump which is not driven by a turbine or may use any other type of pump which has an annular rotor assembly in which the combustion chamber is nested.

As described, the rocket engine 2 features a unique layout with its turbopumps integrated into the engine. The integrated turbopumps are arranged circumferentially around the combustion chamber and there is also an independent turbopump for both the propellants. This unique, integrated turbopump engine architecture is significantly lighter than existing designs and has a significantly lower part count due to the level of integration. Lower mass means less materials and therefore lower costs, as does the lower part count. Furthermore, the manufacturing and assembly time is anticipated to be significantly less than a conventional design. All these factors, plus the choice of hydrolox fuel (the combination of liquid oxygen and liquid hydrogen) is likely to make this the most environmentally sustainable rocket engine of all time.

The rocket engine is able to provide a very competitive specific impulse Isp without using a staged combustion cycle. Engines which use a staged combustion cycle are typically complex, heavy, expensive and difficult to start, and require more exotic materials for the turbopumps to withstand the higher temperatures. Nevertheless, the layout of the present invention may also be beneficially applied to staged combustion cycle engine, such a full-flow staged combustion engine, as well as other engine cycles such as closed expander cycle engines.

As described above, the rocket engine 2 may be based on hydrolox fuel. Hydrolox provides the highest specific impulse of conventional rocket propellants, and it has the significant added benefit in that no carbon dioxide is released when it is burnt.

High chamber pressures in rocket engines lead to higher performing rocket engines and therefore higher specific impulses. However, high chamber pressures require more powerful and therefore heavier turbopump machinery to generate the higher fuel pressures required and therefore lead to higher masses for these high efficiency engines. Whilst certain other liquid propellant choices allow for higher chamber pressure without the engine mass becoming too large, the problem with making high chamber pressure yet lightweight hydrolox rocket engines is fundamentally to do with the extremely low density of liquid hydrogen compared to all other rocket fuels in use. Liquid hydrogen has a very low density compared to most other common liquid fuels for orbital vehicles. For example, liquid hydrogen has a density of 70 kg/m3, whereas RP1 (Kerosene) has a density of 800 to 1000 kg/m3 and liquid methane has a density of 660 kg/m3. Essentially, liquid hydrogen is an order of a magnitude lower density than other rocket fuels which has a huge impact on the engines design and specifically the turbopumps. Generating high fuel pressures with a centrifugal pump is a big challenge for a liquid hydrogen turbopump. To date, hydrolox engines have typically been the lowest thrust-to-weight ratio engines (the best is around 65, compared to conventional propellant engines that have achieved over 200). Consequently, there are currently no economically viable hydrolox only rockets in operation.

The power-to-weight ratio of the rocket engine 2 may be sufficient to allow it to be used solely during the take-off of a rocket without the need for supplemental power from solid rocket boosters and thus without generating any carbon.

It will be appreciated that positional references such as upper, lower, upwards, downwards, top, bottom, etc. are made with respect to the rocket engine being arranged vertically.

The invention is not limited to the embodiments described herein, and may be modified or adapted without departing from the scope of the present invention.

Claims

CLAIMS1. A rocket engine comprising: a combustion chamber; a first pump comprising a rotor assembly for pumping a first propellant to the combustion chamber; wherein at least a portion of the rotor assembly is annular and is arranged so as to encircle and rotate relative to the combustion chamber.
2. A rocket engine as claimed in claim 1, further comprising a second pump comprising a rotor assembly for pumping a second propellant to the combustion chamber.
3. A rocket engine as claimed in claim 2, wherein at least a portion of the rotor assembly of the second pump is annular and is arranged so as to encircle and rotate relative to the combustion chamber.
4. A rocket engine as claimed in claim 2 or 3, wherein the rotor assembly of the first pump overlaps in an axial direction with the rotor assembly of the second pump.
5. A rocket engine as claimed in any one of claims 2 to 4, wherein the first pump and the second pump are arranged in opposite directions.
6. A rocket engine as claimed in any one of the preceding claims, wherein the rotor assembly of the first and/or second pump comprises an impeller; wherein the impeller is annular and is arranged so as to encircle and rotate relative to the combustion chamber.
7. A rocket engine as claimed in any one of the preceding claims, wherein the rotor assembly of the first and/or second pump further comprises a turbine configured to drive the first and/or second pump.
8. A rocket engine as claimed in claim 7, wherein the turbine is annular and is arranged so as to encircle and rotate relative to the combustion chamber.
9. A rocket engine as claimed in any one of the preceding claims, wherein the first pump is fluidically connected to the combustion chamber via a first propellant channel.
10. A rocket engine as claimed in any one of the preceding claims, wherein the first pump is fluidically connected to a chamber cooling channel which surrounds the combustion chamber and extends axially along its length.
11. A rocket engine as claimed in claim 10 when appended to claim 7, wherein the chamber cooling channel is fluidically connected to a turbine drive channel downstream of the combustion chamber; wherein the turbine of the first and/or second pump is disposed within the turbine drive channel.
12. A rocket engine as claimed in claim 11, wherein the turbine drive channel comprises a bypass duct and a bypass control valve which allows the first propellant to bypass the turbine of one or both of the first and second pumps.
13. A rocket engine as claimed in claim 11 or 12, wherein the turbine drive channel is fluidically connected to a turbine exhaust channel downstream of the turbine of the first and/or second pump; wherein the turbine exhaust channel extends alongside the first propellant channel so as to provide heat transfer therebetween.
14. A rocket engine as claimed in claim 13, wherein a portion of the turbine exhaust channel is radially outward of the first propellant channel and a portion of the turbine exhaust channel is radially inward of the first propellant channel; wherein the turbine exhaust channel and the first propellant channel each comprise a plurality of channel portions which are interleaved to allow the turbine exhaust channel to cross over the first propellant channel.
15. A rocket engine as claimed in claim 13 or 14, wherein the turbine exhaust channel terminates within a nozzle of the engine.
16. A rocket engine as claimed in claim 15, wherein the turbine exhaust channel terminates downstream of a throat of the nozzle.
17. A rocket engine as claimed in any one of the preceding claims, wherein said annular portion of the rotor assembly is concentric with the combustion chamber.
18. A rocket engine as claimed in any one of the preceding claims, wherein the first propellant is fuel and second propellant is oxidizer.
19. A rocket engine as claimed in any one of the preceding claims, wherein the first propellant is liquid hydrogen and the second propellant is liquid oxygen.