JPH0196452A - Engine provided with propellant heating section - Google Patents

Abstract

PURPOSE:To improve specific propulsion of propellant and to improve engine performance by arranging a heatexchanger in a propellant flow path for pressure feeding propellant into a combustion chamber and improving enthalpy of propellant and feeding propellant into the combustion chamber. CONSTITUTION:Propellant A is pressure fed through a pump 1 which is driven by a turbine 3 and cools a rocket nozzle 2, then drives the turbine 3 and fed into a combuster 5. Propellant B is simultaneously pressure fed into the combustion chamber 5 through a pump 4 which is driven by same turbine 3 so as to burn the propellant A. Propellant A fed through the pump 1 passes through a heatexchanger 6 arranged in the way where it is heated then fed to the cooling circuit of the nozzle 2. Consequently, enthalpy of propellant is improved so as to improve specific propulsion of the propellant thus improving the engine performance.

Description

DETAILED DESCRIPTION OF THE INVENTION (Field of Industrial Application) The present invention relates to improving the performance of engines that utilize jet flow, and particularly to engines with increased specific impulse, such as rocket engines and ramjet engines.

(Prior Art) Conventionally, in a rocket engine or the like, liquefied fuel and an oxidizer are loaded and simultaneously pumped into a combustion chamber and combusted. In order to improve the performance of a rocket engine, it is necessary to improve the specific impulse, but this specific impulse is determined by the internal energy of the propellant, and in order to increase the specific impulse, it is necessary to increase the specific impulse by using only fuel and oxidizer. Currently, efforts are being made to find the optimal combination of the following. However, the internal energy is uniquely determined by the combination of fuel and oxidizer and cannot be exceeded. Furthermore, since these propellants are loaded in a liquefied state, the internal energy is further reduced by this liquefaction.

(Problems to be solved by this invention) This invention aims to increase the enthalpy of the propellant that is pumped into the combustion chamber, overcome the above-mentioned limitations, and improve the specific impulse to obtain a high-performance engine. It is.

(Means for Solving the Problem) The engine of this invention has a heat exchanger disposed in the flow path of the propellant that is pumped into the combustion chamber to increase the enthalpy of the propellant and supply it to the combustion chamber. It is characterized by

With this configuration, the propellant pumped into the combustion chamber is heated by a heat exchanger placed in its flow path to increase its enthalpy, thereby increasing the specific impulse of the propellant and improving engine performance. can be done.

Any suitable heat source may be used as the heat source on the high temperature side of the heat exchanger, but solar heat can be advantageously used in outer space, and heat generated by aerodynamic heating of the aircraft body can be advantageously used in high-speed flying objects in the air. Furthermore, the heat exchanger is preferably arranged between the propellant pump and its driving turbine.

(Examples) Hereinafter, the present invention will be described in detail by examples with reference to the drawings.

FIG. 1 shows the fuel cycle of a rocket engine implementing this invention. Propellant A is, for example, liquid hydrogen, and propellant B
is, for example, liquid oxygen. In conventional rockets,
The propellant is supplied to the propellant by a pump 1 driven by a turbine 3.
It cools the rocket nozzle 2, drives the turbine 3, and is fed into the combustor 5. On the other hand, the propellant B1 is simultaneously pumped into the combustion chamber 5 by the pump 4, which is also driven by the turbine 3, and the propellant A is combusted.

In the rocket of this invention, the propellant A sent out by the pump 1 passes through a heat exchanger 6 disposed midway, is heated, and then flows into the cooling circuit of the nozzle 2.

In such a rocket engine, let us consider the case where heat is supplied from the heat exchanger by solar heat in space.6 If hydrogen and oxygen are used as propellants and the nozzle opening ratio is set to 200, the heat supply The relationship between the change in specific impulse and the amount is shown in Figure 2.

In the figure, zsp indicates specific impulse, and Po indicates combustion chamber pressure. For example, if 40oOkJ-kg-" of heat is supplied, 8
An increase in specific impulse of ~9 seconds is observed.

Figure 3 shows the results of a trial calculation of how much heat collection area is required if solar heat is to be used to supply this amount of heat. AP in the figure indicates the heat collection area.

As mentioned above, when the amount of heat supplied is 4000 kJ-kg-'', the combustion chamber pressure PC = 5 MPa requires a heat collection area of about 65 Or/, but a small rocket with PC = IMPa requires a heat collection area of about 130 This can be achieved with the heat collection area.

FIG. 4 shows the change in specific thrust crab 3P when the hydrogen/oxygen mixture ratio is changed. The combustion conditions are combustion chamber pressure P c
= IMPa, nozzle opening ratio 200. It can be seen that as the amount of heat supplied increases, the mixture ratio exhibiting the maximum specific impulse moves away from chemical equivalence, and the efficiency of the propellant is improved compared to when the propellant is combusted only.

Furthermore, in the engine shown as this example,
Another effect is that the pressure throughout the cycle is lower than before. That is, heating the propellant increases the enthalpy level of the propellant. Therefore, in topping cycles such as expander cycles, the propellant temperature after regeneration cooling, that is, the turbine inlet temperature, increases.
The pressure ratio required to drive the turbine is smaller than before.

For this reason, the maximum cycle pressure, which appears as the pressure at the outlet of the propellant pump on the turbine drive side, is lower than before, and the pressure of the entire cycle is also lower than before.

Figure 5 illustrates this effect. That is, in the engine shown in FIG. 1, hydrogen is used as the driving fluid for the turbine 3, and the change in the outlet pressure Pfu of the pump 1, which is the highest pressure in the cycle, with respect to the amount of heat supplied is shown as ν). Combustion chamber pressure P. When Pc is small, this effect is not significant, but as the combustion chamber pressure Pc becomes large, as is clear from the figure, there is a noticeable pressure drop. The results are shown here based on calculations using liquid oxygen as the propellant, but of course other propellants can produce similar effects, and the source of heat is not limited to solar heat, but in supersonic aircraft, etc., there is heat from aerodynamic heating of the aircraft body, etc. , an appropriate heat source can be used. Furthermore, although the embodiment has been described with the heat exchanger disposed in the flow path of liquid hydrogen used to drive the turbine, heating the oxidizing agent has the same effect of improving the enthalpy of the propellant.

(Effects of the Invention) As described above, the rocket engine of the present invention has a simple configuration in which a heat exchanger is disposed for supplying heat to the propellant passage, and the rocket engine exceeds the inherent limit of propellant. This not only improves thrust, but also reduces fuel cycle pressure and can contribute to reducing the weight of the engine structure.

Note that this technology is applicable not only to rockets but also to scramjets, ramjets, etc.

[Brief explanation of the drawing]

Fig. 1 is a conceptual diagram of the operating system showing one embodiment of the fuel cycle of a rocket engine implementing this invention, Fig. 2 is a graph showing the relationship between the amount of heat supplied to the propellant and specific impulse, and Fig. 3 is the propulsion system. A graph showing the amount of heat supplied to the propellant and the required solar heat collection area, Figure 4 is a graph showing the change in specific impulse TSP when the hydrogen/oxygen mixture ratio is changed, and Figure 5 is a graph showing the change in specific impulse TSP for the propellant. This is a graph showing the relationship between heat supply h (and the maximum pressure of the cycle, and each symbol in the figure is 1.
．． 4: Pump 2: Rocket nozzle 3: Turbine 5: Combustor 6: Heat exchanger. Fig. 1 A B Fig. 2 Supplied heat amount kJ-kg-1 Fig. 3 Supplied heat amount kJ-kg-' Fig. 4 Hydrogen/oxygen

Claims (1)

[Scope of Claims] 1) A propellant characterized in that a heat exchanger is disposed in the flow path of the propellant that is pumped into the combustion chamber to increase the enthalpy of the propellant so that it is supplied to the combustion chamber. 2) An engine having a heating part 2) An engine having a propellant heating part according to claim 1, wherein the high temperature side of the heat exchanger is supplied with solar heat 3) The heat exchanger 4) An engine having a propellant heating section according to claim 1, wherein the heat exchanger is disposed in a propellant flow path for driving a turbine for a propellant pressure pump. , an engine having a propellant heating section according to claim 1, wherein the propellant heating section is disposed between a propellant pressure-feeding pump and its driving turbine.