WO2011107090A2

WO2011107090A2 - Method for operating a rocket engine and rocket engine

Info

Publication number: WO2011107090A2
Application number: PCT/DE2011/075019
Authority: WO
Inventors: Olaf Przybilski
Original assignee: Technische Universität Dresden
Priority date: 2010-03-05
Filing date: 2011-02-10
Publication date: 2011-09-09
Also published as: DE102010010265B4; WO2011107090A3; DE102010010265A1

Abstract

The invention relates to a method for operating a rocket engine, which method makes it possible to provide single-stage carrier rockets that achieve an optimal thrust and a maximum specific impulse over the combustion duration of the engine. The invention further relates to a rocket engine that allows the method according to the invention to be carried out. In the method according to the invention, at least one hydrocarbon is reacted with oxygen in a first combustion chamber of the rocket engine and hydrogen is reacted with oxygen in a second combustion chamber of the rocket engine. The reaction in the first combustion chamber is maintained until the hydrocarbon is consumed; the reaction in the second combustion chamber is continued until the hydrogen is consumed. The fuel amounts are selected so that the hydrocarbon is consumed earlier than the hydrogen. The rocket engine, designed as a plug nozzle engine, for carrying out the method comprises a main body (1), at the lower end of which a central cone (2) (plug nozzle/spike) is connected, wherein the main body (1) is surrounded by a first annular combustion chamber (3.1), which has a first annular throat nozzle (6.1), and a second annular combustion chamber (3.2) having a second annular throat nozzle (6.2) is located in the central cone (2).

Description

Method of operating a rocket engine and rocket engine

The invention relates to a method of operating a rocket engine, which makes it possible to provide single-stage launch vehicles which achieve optimum thrust and maximum specific momentum throughout the duration of the engine's combustion. Furthermore, the invention relates to a rocket engine, which enables the implementation of the method according to the invention.

The flight of a rocket must take into account sometimes conflicting requirements and changing environmental conditions in order to obtain an effective rocket engine.

First of all, when the rocket is launched and immediately afterwards, a "heavy" rocket fuel is needed, which provides a great boost to its combustion, and as the climb progresses, a "lighter" fuel is needed to provide a high specific impulse.

Current carrier systems meet these requirements by being multi-level. A distinction is made between tandem and parallel-tiered carrier systems.

In tandem staged carrier systems several individual stages are stacked on top of each other, which use different fuels during combustion. In the first stage, a heavy fuel is burned, which provides a great boost. In the next stage (s), fuel is easily combusted, providing a high specific impulse. After consuming the fuel, the associated stage is disconnected and the rocket engine of the next stage activated. The disadvantage is that always dead payload in the form of non-operating rocket motors must be carried and accelerated.

In parallel-staged carrier systems, such as the Ariane 5, the individual stages are arranged in parallel and activated at the same time, however, the parallel stages that work with heavy fuel (booster), operated only in the start phase and then discarded. It is advantageous that the installed engine plant load is fully utilized during the entire drive time; In addition, no engine must be carried as a dead payload during the drive phase.

The boosters are very often operated on solid fuels, but have some significant disadvantages. Solid boosters must always be transported in a loaded condition, which is a disadvantage in terms of safety (explosion risk) and effort (transport weight) during transport compared to liquid fuel rockets, which are usually refueled just before take-off. Furthermore, the combustion products of the solid rocket are usually more environmentally harmful than the liquid fuels. In addition, the boosters are hardly controllable in the thrust regime and can not be switched off after ignition.

Furthermore, the usually parabolic or bell-shaped expansion nozzles of the respective engines are adapted only to the air pressure of a narrow height range, since they work optimally only in the adapted regime. This results from the condition that the pressure of the outflowing combustion gases at the nozzle end must correspond to the pressure of the outside atmosphere, so that the engine can develop its full power. Since the gas pressure at the nozzle end remains approximately constant over the entire burning time, but the pressure of the atmosphere decreases continuously with increasing distance to the earth's surface, such engines only work optimally for a short time.

From the prior art solutions are known in which the engine geometry is changed to adapt to the changing ambient pressure. Such solutions, however, could not prevail in practice, since they are elaborate production and control technically complex and also mass-intensive.

Active rocket engine height adjustment is not necessary with Aerospike or Plug Nozzle engines. Engines of this type theoretically achieve 90% efficiency regardless of air pressure. They surpass any traditional bell engine.

With the known plug-Nozzle engines, as described for example in US 2005 022 36 95 A1, but not the problem is solved that one for the Different phases of the rocket flight further require stepped carrier systems, which provide first for a high thrust and later for a high specific impulse.

To counter this, attempts have been made to develop aerospike or plug-in engines, which after the launch of the rocket, first with a heavy fuel, which provides a great boost, and then with a light fuel, which for provides a high specific impulse to operate.

However, developments in this direction have so far failed u. a. The fact that the engines only work effectively when the mass flow through the nozzles is adapted to the nozzle cross-section. When launching the rocket, a very high mass flow through the engine is needed. As soon as the rocket has lost much of its mass due to fuel consumption, a lower mass flow through the engine would be sufficient to further accelerate the rocket. If one reduces the mass flow in such an engine, without adapting the nozzle cross-section to this reduced mass flow, the generated specific impulse is no longer sufficient to accelerate the rocket to its final speed.

The object of the invention is to provide a method for operating a rocket engine for a single-stage rocket, which meets the changing requirements in terms of thrust and specific momentum in the flight of a single-stage launch vehicle. Furthermore, a simple design plug-Nozzle engine low mass to be provided for performing the method.

The object is achieved by the features of the independent claims 1 and 6, advantageous embodiments of the invention can be found in the dependent claims.

According to the invention, the object is achieved in that, depending on the altitude of the rocket, the thrust and the specific momentum are varied in that differing and controllable in their quantity fuels are reacted with an oxidizer in the rocket engine. At the start At least one hydrocarbon with oxygen and in a second combustion chamber of the rocket engine hydrogen are reacted with oxygen of the rocket in a first Brennkannnner the rocket engine. The reaction in the first combustor is maintained until the hydrocarbon is exhausted. In the second combustion chamber, which is smaller than the first, hydrogen and oxygen are injected until the hydrogen is consumed. The amounts of fuel are such that the hydrocarbon is consumed earlier than the hydrogen.

Prior to injection of the fuels into the combustors, they are routed along the walls of the combustors, removing heat energy from the walls and thus cooling them. As a result, the cooling of the combustion chambers is guaranteed from the start to the end of the combustion.

Preferably, only one hydrocarbon is burned in the first combustion chamber and hydrogen is burned in the second combustion chamber. The combustion of the heavy hydrocarbon - preferably ethanol or methane - when the rocket is launched, causes the engine to develop a very high thrust. Once the heavy fuel is consumed, the engine is powered only with the light fuel. Since the total mass of the rocket is greatly reduced after consumption of the hydrocarbon, a correspondingly lower engine power is required; the thrust produced by the combustion of hydrogen in the second combustion chamber is sufficient for this. At the same time, the required high specific impulse is generated in order to bring the rocket to its final speed.

In a further embodiment variant of the method, provision is made for hydrogen to be burned in the first combustion chamber in addition to the hydrocarbon; in the second combustion chamber, only hydrogen is reacted with oxygen.

The use of liquid ethanol or methane as a heavy fuel instead of solid fuel is advantageous because the fuel is controllable for combustion, si- to handle more easily and its combustion is much more environmentally friendly than the solid fuels.

The rocket engine for carrying out the method according to the invention is designed as a plug-Nozzle engine. It consists of a main body, at the lower end of which a central cone (plug-nozzle or spike) connects. The main body is surrounded by a first annular combustion chamber, which opens in its lower end in a Ringhalsdüse. In the central cone there is a second annular combustion chamber with a second annular neck nozzle, both of which are smaller than the first combustion chamber and annular neck nozzle. The ring-neck nozzles of both combustors are oriented so that the flow vectors of the combustion gases passing therethrough are parallel to the surface of the central cone.

About injectors, which are located on the opposite sides of the ring neck nozzles of the annular combustion chambers, the fuels (hydrocarbon, hydrogen and oxygen) are injected into the combustion chambers and reacted there. The injectors of the first combustion chamber may be tricaxial swirl nozzles, which allow to simultaneously inject two fuels (eg, ethanol and hydrogen) and oxygen.

In the preferred embodiment of the rocket engine, ethanol or methane is supplied to the engine in the liquid state and initially held in two annular tubes, the ring collectors. The first (inner) ring collector is arranged concentrically within the main body, the second (outer) ring collector is located on the outside of the upper ring combustion chamber at the level of the ring neck nozzle.

From the inner ring collector forth the fuel is guided on the outside of the inner wall of the annular combustion chamber on the one hand to the injection nozzles and the other to an annular cooling nozzle below the Ringhalsdüse. From the outer ring collector forth the fuel flows along the outside of the outer wall of the annular combustion chamber to the injection nozzles. When flowing along the hot walls of the annular combustion chamber, the fuel extracts heat energy from the walls. As a result, the walls of the annular combustion chamber are cooled and protected against overheating and destruction by the hot combustion gases.

In the transition region from the main body to the central cone is located at the level of Ringhalsdüse a turbine exhaust ring collector, in which the comparatively cold exhaust gases of the fuel pump turbines are first collected and then fed to the cooling nozzle, where they emerge together with liquid fuel from this and hereinafter at the Flow along the surface of the central cone and burn at a low temperature.

The turbine exhaust gases and the fuels flowing along the surface of the central cone perceive two functions; they cool the central cone and act as an insulating layer between the central cone and the hot combustion gases exiting through the ring neck nozzle.

Furthermore, at the level of the annular neck nozzle of the second annular combustion chamber, there are two further hydrogen ring collectors, the first annular collector being enclosed by the annular neck nozzle and the second enclosing the annular neck nozzle. From the ring neck nozzles, the hydrogen is guided along the walls of the second annular combustion chamber, wherein the hydrogen passes from the liquid to the gaseous state. The gaseous hydrogen is finally fed to the injection nozzles. A portion of the hydrogen retained in the annular collector enclosed by the annular neck nozzle flows out through a second cooling nozzle located immediately below the annular ring nozzle of the second annular combustion chamber. The outflowing hydrogen adapts to the surface of the central cone and thus again forms an insulating layer between the hot combustion gases and the central cone.

It is envisaged to discharge a portion of the turbine exhaust gases through a straightening nozzle in the tip of the central cone, which provide for a homogenization of the flow direction of the converging at the top of the central cone combustion gases. Thus, an extremely effective rocket propulsion system is provided which operates with high efficiency throughout the entire rocket-traversed altitude range, while accommodating the different requirements of thrust and specific momentum over the rocket ascent phase.

The invention will be explained in more detail below with reference to an embodiment and Figures 1 and 2; show here:

Fig. 1 is a rotationally symmetrical engine in cross section

Fig. 2 shows the lower portion of the engine shown in Fig. 1 in an enlarged view.

The engine shown in Fig. 1 consists of a cylindrical main body 1 and a subsequent central cone 2 (plug-nozzle or spike) with an opening angle of about 60 °. The main body 1 is enclosed by the first annular combustion chamber 3.1, which has an inner wall 4 and an outer wall 5. In the area of the transition from the main body 1 to the central cone 2, the first annular neck nozzle 6.1 is located. Within the first annular combustion chamber 3.1 of the annular neck nozzle 6.1 opposite the injectors 7.1 are introduced.

Within the main body 1 and coaxial with its body axis is the first, inner hydrocarbon ring collector 8; outside the first annular combustion chamber 3.1 is the second, outer hydrocarbon ring collector 9. Both hydrocarbon ring collector 8; 9 are applied during operation of the engine with heavy fuel from the ring collectors 8; 9, starting along the walls 4; 5 of the annular combustion chamber 3.1 is passed to the injectors 7.1 of the first annular combustion chamber 3.1.

The fuel (ethanol) is injected into the first combustion chamber 3.1 together with the oxidizer (liquid oxygen) via the injection nozzles 7.1. There, the individual fuel components react with each other. The resulting combustion gases 10 flow through the first ring neck nozzle 6.1 from the annular combustion chamber 3.1. The opening of the Ringhalsdüse 3.1 is aligned orthogonal to the surface of the central cone 2, so that the combustion gases 10 flow substantially parallel to the surface of the central cone 2 along. At the level of the first annular neck nozzle 6.1 is in the transition region of the main body 1 to the central cone 2 of the turbine exhaust ring collector 1 1. In it, the comparatively cold exhaust gases of the pump turbine are collected and then fed to the cooling nozzle 12.1. The cooling nozzle 12.1 is also supplied with a part of the fuel from the inner hydrocarbon ring collector 8, which was guided along the lower part of the inner wall 4 of the first annular combustion chamber 3.1. Like the ring neck nozzles 6.1; 6.2, the cooling nozzle 12.1 is aligned such that the turbine exhaust gas hydrogen mixture 13 flows parallel to the surface of the central cone 2. The cooling nozzle 12.1 is located below the first annular neck nozzle 6.1, so that the turbine exhaust gas fuel mixture 13 forms a cooling film between the central cone 2 and the hot combustion exhaust gases 10.

Within the central cone 2, but below the engine parts described above, there is the second annular combustion chamber 3.2 with a Ringhalsdüse 6.2, at the height of the hydrogen ring collector 14.1 and 14.2 are located. These are supplied with hydrogen, which is guided along the walls of the annular combustion chamber 3.2 to the injection nozzles 7.2. Like the hydrocarbon in the ring neck nozzle 3.1 here, the hydrogen absorbs heat from the walls of the annular combustion chamber 3.2 and cools it with it. In the hydrogen ring collectors 14.1; 14.2 the hydrogen is still liquid at first; he goes into the gaseous state of matter while he flows along the walls of the annular combustion chamber 3.2. The gaseous hydrogen is injected together with liquid oxygen into the annular combustion chamber 3.2 and reacted.

The resulting during combustion combustion gases 10 are ejected through the Ringhalsdüse 6.2 and continue to flow parallel to the surface of the central cone 2. To protect the central cone 2 from the hot combustion gases 10, it is provided that a cooling nozzle 12.2, below the annular combustion chamber 6.2 sits to expel pure hydrogen, which serves as an insulating layer between the combustion gases 10 and the central cone 2.

In order to reduce turbulence at the top of the central cone 15 turbine exhaust gases 16 are blown through the directional nozzle. List of reference numbers used

1 main body

2 central cone

3.1 first annular combustion chamber

3.2 second annular combustion chamber

4 inner wall

5 outer wall

6.1 first ring neck nozzle

6.2 second ring neck nozzle

7.1 Injector

7.2 Injector

8 inner hydrocarbon ring collector

9 outer hydrocarbon ring collector

10 combustion gases

1 1 Turbine exhaust ring collector

12.1 first cooling nozzle

12.2 second cooling nozzle

13 turbine exhaust fuel mixture

14.1 first hydrogen ring collector

14.2 second hydrogen ring collector

15 straightening nozzle

16 turbine exhaust

Claims

claims

A method of operating a rocket engine, in which, depending on the altitude of the rocket, the thrust and the specific momentum are varied by reacting differing and controllable fuels with an oxidizer in the rocket engine, characterized in that when the rocket is launched in a first combustor of the rocket engine, at least one hydrocarbon is reacted with oxygen and in a second combustor of the rocket engine hydrogen is reacted with oxygen, the reaction in the first combustor is maintained until the hydrocarbon is exhausted, and the Reaction in the second combustion chamber is continued until the hydrogen is consumed, wherein the amounts of fuel are so dimensioned that the hydrocarbon is consumed earlier than the hydrogen.

2. A method for operating a rocket engine according to claim 1, characterized in that in the first combustion chamber, both hydrocarbon and hydrogen are reacted with oxygen.

3 A method for operating a rocket engine according to any one of claims 1 and 2, characterized in that the reactions in the first and the second combustion chamber are started simultaneously.

4. A method for operating a rocket engine according to one of claims 1 to 3, characterized in that the hydrocarbon prior to injection into the first combustion chambers (6.1) on the walls (4; 5) and / or the hydrogen prior to injection into the combustion chamber 3.2 is guided along the walls of the combustion chamber 3.2 along.

A method for operating a rocket engine according to any one of claims 1 to 4, characterized in that the hydrocarbon is ethanol or methane. Rocket engine, designed as a plug-nozzle engine, for carrying out the method according to one of claims 1 to 5, consisting of a main body (1), at the lower end of which a central cone (2) (plug-nozzle / spike) connects, characterized in that the main body (1) is enclosed by a first annular combustion chamber (3.1) having a first annular neck nozzle (6.1) and a second annular combustion chamber (3.2) with a second annular neck nozzle (6.2) is present in the central cone (2) the ring neck nozzles (6.1, 6.2) are aligned so that the flow vectors of the combustion gases passing through them are parallel to the surface of the central cone (2). Rocket engine according to claim 6, characterized in that at the walls of the annular combustion chambers (3.1, 3.2) opposite the ring-necked nozzles (6.1, 6.2) are injection nozzles (7.1, 7.2). Rocket engine according to claim 7, characterized in that the injection nozzles (7.1) on the ring neck nozzle (6.1) opposite wall of the annular combustion chamber (3.1) tricaxial spin nozzles. Rocket engine according to one of claims 6 to 8, characterized in that in the region of the transition of the main body (1) to the central cone (2) a cooling nozzle (12.1) is located, through the central cone (2) and via the Ringhalsdüse (6.1) extended inner wall (4) is limited. A rocket engine according to any one of claims 6 to 9, characterized in that inside the main body (1) there is an inner hydrocarbon ring collector (8) for hydrocarbons whose axis is coaxial with that of the main body connected to the cooling nozzle (12.1) and with the injection nozzles (7.1) is in communication. Rocket engine according to one of claims 6 to 10, characterized in that outside the annular combustion chamber (3.1) at the level of the annular neck nozzle (6.1) an outer hydrocarbon ring collector (9) for coal which is in communication with the injection nozzles (7.1). Rocket engine according to one of claims 6 to 1 1, characterized in that within the central cone (2) a turbine exhaust ring collector (1 1) is located, the body axes are coaxial with each other and is in communication with the cooling nozzle (12.1). Rocket engine according to one of claims 6 to 12, characterized in that within the central cone (2) there are a first hydrogen ring collector (14.1) and a second hydrogen ring collector (14.2) which are in communication with the injection nozzles (7.2). Rocket engine according to claim 13, characterized in that the first hydrogen ring collector with a cooling nozzle (12.2) is in communication, which is arranged in the central cone (2) below the annular neck nozzle 6.2. Rocket engine according to one of claims 6 to 14, characterized in that in the tip of the central cone (2) a directional nozzle (15) is arranged.