JPH0452859B2 - - Google Patents

Description

DETAILED DESCRIPTION OF THE INVENTION The invention relates to a method according to the preamble of claim 1 for operating a liquid rocket propulsion mechanism and to a rocket propulsion mechanism for carrying out this method.

According to Patent Publication No. 2743983 of the Federal Republic of Germany,
Liquid rocket propulsion mechanisms working in a two-liquid manner are known, in which the heated propulsion nozzle walls and combustion chamber walls are cooled and heated and evaporated. A portion of the hydrogen is used to drive the pump drive turbine. The working substances still contained in the turbine exhaust gas are utilized in the subsequent secondary propulsion nozzle.

A fundamental disadvantage of this method is the relatively small amount of energy supplied in terms of hydrogen during combustion chamber cooling and propulsion nozzle wall cooling. Furthermore, since the turbine output is not sufficient to be multiplied, the upper limit of the combustion chamber pressure that can be achieved is limited, and currently it is approximately
50 bar.

This disadvantage is due to the fact that in classic two-propellant rocket propulsion systems, the power required for one or more pump-driven turbines is not achieved by a fraction of the rocket propellant in the secondary fuel stream. This was avoided by generation with an energized auxiliary gas generator. However, in this method, the propellant portion that is branched off to operate the auxiliary gas generator under the desired high and extremely high pressures in which the propulsion combustion chamber is located is subject to the propulsion power loss associated with this branching. quickly becomes so large that it exceeds the gain due to the higher pressure in the propulsion combustion chamber, so even with this modified process the overall efficiency remains constant even at a constant high or constant combustion chamber pressure. The pressure drops again from a high level. Furthermore, in view of the permissible operating temperatures of the turbine blades, it is not possible to operate the auxiliary combustion chamber or the secondary liquid flow combustion chamber with optimal efficiency, that is to say stoichiometrically. As a result, a significant portion of the working material within the propellant is lost to the outside through secondary liquid circulation.

The above-mentioned drawbacks are a result of so-called rocket mainstream methods, such as those disclosed, for example, in the above-mentioned German patent publication.
In this case, a pre-combustion chamber is connected in front of the main combustion chamber in response to the fuel flow, and in this pre-combustion chamber, for example, all the hydrogen and A portion of the oxygen is reacted, thereby generating a temperature at which the drive gas containing excess hydrogen can still be used for the subsequent pump drive turbine in response to the fuel flow. This is avoided by The turbine exhaust gas then flows into the next main combustion chamber, into which residual oxygen is introduced for stoichiometric combustion.

A certain drawback of the mainstream process principle is that the injection head of the main combustion chamber is passed by relatively still very hot turbine exhaust gas, which results from the entire amount of hydrogen participating in the reaction process, in addition to a partial stream of oxygen. This thermal load makes construction of the injection head extremely difficult and expensive. Furthermore, in addition to the partial amount of oxygen, the entire amount of hydrogen must be increased to a high injection pressure, which also requires a high pump power.

It is therefore an object of the present invention to create an operating method for a liquid rocket propulsion mechanism which eliminates the disadvantages of hitherto known propulsion devices while at the same time retaining their advantages, and which operates with higher efficiency, and which The aim is to create favorable structural conditions for the injection head of the engine combustion chamber.

This problem is achieved according to the invention in a rocket propulsion mechanism of the type described at the outset as follows. That is, the propellant, in particular hydrogen, which has been previously heated in the propulsion nozzle walls and the combustion chamber walls, is required for the required pump drive power before it enters the turbine or turbines as a turbine drive gas. heat is supplied by a heat exchanger loaded with combustion gases generated in an auxiliary combustion chamber operating stoichiometrically from the rocket propellant or from partial amounts of hydrogen and oxygen, and the exhaust gas of said auxiliary combustion chamber is is introduced into the region of the propulsion nozzle at a level of pressure lower than the pressure of this exhaust gas.

Within the framework of the invention, the auxiliary combustion chamber is operated with optimum efficiency for generating the combustion gases that operate the heat exchanger, i.e. with a mixing ratio that yields optimum efficiency and in a stoichiometric manner. It is possible to heat the turbine drive gas to such an extent that the required power of the propellant feed pump is achieved and to provide the turbine drive gas with operating capacity. In this case, without incurring power losses, the temperature of the turbine exhaust gas is lowered by the classical two-component method, since the temperature of the turbine drive gas is lower than that of the turbine drive gas. Considering the power of the secondary flow combustion chamber in relation to the temperature of the resulting maximum turbine temperature can be kept lower than in one. The low temperature of the turbine exhaust gas in the method according to the invention is advantageous for the injection head of the propulsion combustion chamber. This is because loading the injection head with heat, which would have any adverse consequences on the combustion chamber of the propulsion mechanism, is thereby avoided.

A further optimization of efficiency and power is achieved according to the invention by the turbine exhaust, which releases heat here on the one hand, before the heat exchanger (with an auxiliary combustion chamber), which is provided upstream of the pump-driven turbine. This is achieved by connecting an additional heat exchanger through which the gas and the turbine drive gas or hydrogen, which absorbs heat, flow through.

A further method step according to the invention makes it possible to increase the power level of the turbine drive and the pump drive without having negative consequences for the injection head of the combustion chamber of the propulsion mechanism. In other words,
The increase in temperature achieved by the additional heat exchanger and thus the decrease in the output of the turbine - which, on the other hand, leads to an increase in the temperature of the turbine exhaust gas - causes the subsequent injection to react immediately to the flow. There are no adverse effects of heat on the head.
This is because the relatively high temperature of the turbine exhaust gas is then lowered again in the heat exchanger additionally provided according to the invention before the heat reaches the injection head. The net result is that a relatively high overall efficiency and an effective power increase are possible within the framework of the invention.

The invention will be explained in detail below with reference to embodiments illustrated in the accompanying drawings.

As is clear from FIG. 1, the entire rocket propulsion mechanism essentially consists of a propulsion combustion chamber 1 comprising an injection head 2 and a propulsion nozzle 3 connected in a convergent-divergent manner, a propellant combustion chamber 1 containing liquid hydrogen H; a storage container 4 for liquid oxygen O, a storage container 5 for liquid oxygen O, a hydrogen feed pump 6, an oxygen feed pump 7, a turbine 8 for driving the feed pump 6, and a turbine 8 for driving the feed pump 7. It consists of a turbine 9 and an auxiliary combustion chamber 10 with an integrated heat exchanger 11.

From the hydrogen feed pump 6, a feed conduit 12 allows hydrogen H to flow through the walls of the propulsion nozzle and the wall of the combustion chamber 1 for cooling purposes. At this time, hydrogen H is heated. A portion of the hydrogen H, referenced Hh, flows via branch line 13 into auxiliary combustion chamber 10 and is introduced into this auxiliary combustion chamber. A supply conduit 14 from the oxygen supply pump 7
opens into the injection nozzle 2, from which a branch line 15 runs into the auxiliary combustion chamber 10. Inside this combustion chamber there is also a part of oxygen O.
The one indicated by Oh is introduced. The combustion gases B generated stoichiometrically in the auxiliary combustion chamber 10 load on the one hand a heat exchanger 11 and on the other hand they are heated for the first time in the injection nozzle wall and the combustion chamber wall. The hydrogen Hw that is present flows through.

Hydrogen Hw passes through the intermediate supply conduit 16 to the heat exchanger 1
1. The hydrogen further heated in the heat exchanger 10 loads both turbines 8 and 9 as a turbine drive gas Ht, where the propellant feed pumps 8 and 9
Generates output to drive. The turbine exhaust gas He is supplied to the injection head 2 via a coupling conduit 17. The exhaust gas AB of the auxiliary combustion chamber is introduced into the propulsion nozzle 3 in a region where the pressure level of this propulsion nozzle is below the pressure value of the exhaust gas AB.

The embodiment shown in FIG. 2 differs from the embodiment shown in FIG. 1 in the following respects. That is, in this embodiment both conduits 16 and 17 serve as an auxiliary heat exchanger 18.
This heat exchanger has the function of supplying heat for the first time to the hydrogen Hw that is heated at the combustion chamber wall and for leaving there, as well as heat for releasing the turbine exhaust gas He′. They differ in that they do Hydrogen Hw', which has already been heated twice, is supplied with heat a third time in the next heat exchanger 10. The extremely hot hydrogen leaving this heat exchanger 10 is then used as a turbine drive gas
Forms Ht′.

[Brief explanation of the drawing]

1 and 2 illustrate two embodiments of a method according to the invention and a rocket propulsion mechanism for carrying out the method. The symbols in the figure are 3... Propulsion nozzle, 8, 9... Turbine, 10... Auxiliary combustion chamber, 11... Heat exchanger, AB...
Exhaust gas, B...combustion gas, Ht...turbine driving gas, Hh, Hw...hydrogen, Oh...oxygen.

Claims (1)

Claims: 1. A propulsion nozzle of essentially convergent-divergent configuration, a propellant feed pump, heated and vaporized by precooling the propulsion nozzle wall and the combustion chamber wall. Rocket propulsion operated by liquid propellants, in particular hydrogen and oxygen, consisting of one or more turbines driven by one of the two propellants, in particular hydrogen, and operating said propellant feed pumps. In the method for operating the mechanism, the propellant, in particular hydrogen Hw, is heated beforehand in the propulsion nozzle wall and in the combustion chamber wall, which flows into one or more turbines 8, 9 as turbine drive gas Ht. The heat required for the previously required pump drive power is loaded with combustion gases B generated in an auxiliary combustion chamber 10 operating stoichiometrically from the rocket propellant or from partial quantities of hydrogen Hh and oxygen Oh. supplied by a heat exchanger 11,
A method for operating the liquid rocket propulsion mechanism, characterized in that the exhaust gas AB of the auxiliary combustion chamber is introduced into the region of the propulsion nozzle 3 at a pressure level lower than the pressure of this exhaust gas. 2. A propulsion nozzle of essentially convergent-divergent configuration, a propellant feed pump, both propellants heated and vaporized by precooling the propulsion nozzle wall and the combustion chamber wall. for operating a rocket propulsion mechanism operated by liquid propellants, in particular hydrogen and oxygen, comprising one or more turbines, in particular driven by hydrogen, and operating said propellant feed pumps; A method in which the propellant, in particular hydrogen Hw, which has been previously heated in the propulsion nozzle walls and the combustion chamber walls, is injected with the necessary amount of hydrogen before it enters one or more turbines 8, 9 as turbine drive gas Ht. A heat exchanger in which the heat required for the pump drive output is loaded with combustion gases B generated in an auxiliary combustion chamber 10 operating stoichiometrically from the rocket propellant or from partial amounts of hydrogen Hh and oxygen Oh. supplied by a vessel 11;
and in a rocket propulsion mechanism for implementing the method for operating the liquid rocket propulsion mechanism, which introduces the exhaust gas AB of the auxiliary combustion chamber into the region of the propulsion nozzle 3 at a pressure level lower than the pressure of this exhaust gas. , heat exchanger 11 and auxiliary combustion chamber 1
0 forms one structural unit. 3 Upstream of the heat exchanger 11, which is arranged upstream of the pump-driven turbines 8, 9 and is equipped with an auxiliary combustion chamber 10, on the one hand the turbine exhaust gas He', which releases the heat here, and on the other hand, the turbine exhaust gas He' which releases the heat here, According to claim 2, an additional heat exchanger 18 is provided, through which the turbine drive gas or hydrogen Hw is absorbed and which has already passed through the propulsion nozzle wall and the combustion chamber wall. rocket propulsion mechanism.