JPH01164700A - Airframe cooling cycle - Google Patents

Abstract

PURPOSE: To improve efficiency by cooling an airframe with a propellant, and spouting the heated propellant after this heat exchange from a nozzle in a missile exposed to a high temperature by aerodynamic heating, etc. CONSTITUTION: The hydrogen fed from a liquid hydrogen tank 2 provided in an airframe 1 exposed to a high temperature is boosted by a pump 3, then it is used for the heat exchange (cooling) with the airframe 1 by a heat exchanger 4. The heat-exchanged hydrogen drives a turbine 6' with its whole quantity, for example, then the hydrogen is exhausted through a nozzle 5. Power is obtained by a large turbine flow in this cycle, the hydrogen temperature at a turbine outlet is relatively high in a range that the pressure ratio required for the turbine 6' is near 1, and a high specific impulse can be expected.

Description

Detailed Description of the Invention (Industrial Application Field) The present invention relates to a cooling cycle for the airframe of a high-speed flying object where the airframe is exposed to high temperatures, and particularly to a cooling cycle in which the energy absorbed by the coolant is used to propel the airframe. .

(Prior art) When a space shuttle re-enters the atmosphere, the aircraft is exposed to high temperatures of up to 2000 degrees Celsius due to aerodynamic heating, etc., and in order to protect the aircraft from this high temperature, measures such as placing heat insulating tiles on the surface of the aircraft have been taken. What is being taught is well known.

(Problem to be solved by this invention) In the case of conventional space shuttle re-entry, etc., the period of exposure to high temperatures is not very long, and the insulating layer prevents the aircraft from being directly exposed to high temperatures. However, in the case of supersonic intercontinental vehicles, etc., the time spent undergoing aerodynamic heating is long, making it difficult to sufficiently prevent the temperature of the vehicle from rising with just the insulation layer.

Furthermore, the protection of the aircraft with a heat insulating layer merely blocked the generated heat and allowed it to escape, but did not actively utilize it.

(Means for Solving the Problem) In this invention, in high-speed flying vehicles such as space planes, the aircraft body is exposed to high-temperature air due to aerodynamic heating, etc. In order to protect the aircraft body from this high temperature, liquid It is cooled by a propellant such as hydrogen, and the heated coolant after this heat exchange is ejected from a nozzle and used for the auxiliary propulsion system.

In order to perform heat exchange, it is necessary to increase the pressure to compensate for the pressure loss caused by heat exchange and pump the coolant, but it is better to use the coolant that has been used to cool the aircraft to open the pump. .

(Example) Examples of this invention will be shown below.

Figure 1 shows a conceptual diagram of the airframe cooling cycle when fuel is used as the coolant. In this example, hydrogen was used as the fuel. In the figure, numeral 1 is an aircraft body exposed to high temperatures, and the frame above it is a conceptual diagram of a cooling system installed within the aircraft body.

Hydrogen sent from the liquid hydrogen tank 2 is pressurized by a pump 3 and then used for heat exchange (cooling) with the aircraft body in a heat exchange section 4.

The hydrogen after heat exchange is exhausted through the nozzle 5, and propulsive force is obtained. In the figure, 6 indicates an opening device for the pump 3, but any device may be used for this drive.

FIG. 2 shows an embodiment in which this pump housing system is a turbine. Figure 2 (a) shows the total amount of hydrogen at the turbine 6 after heat exchange.
' is opened, and then hydrogen is exhausted through the nozzle. In FIG. 2B, part of the hydrogen after heat exchange is used to drive the turbine 6', and the rest is exhausted through the nozzle 5.

Furthermore, the hydrogen used to drive the turbine is exhausted from the nozzle 5'. In the figure, the same components as in FIG. 1 are represented by the same symbols.

The cycle (a) is a type that obtains power with a large turbine flow rate. In a range where the pressure ratio required by the turbine is close to 1, the hydrogen temperature at the turbine outlet is also relatively high, and high specific impulse is expected.

The cycle (b) is a type that obtains power at a relatively large turbine pressure ratio. For this reason, the hydrogen temperature at the turbine outlet drops considerably, and high specific impulse cannot be expected from the turbine exhaust hydrogen from the nozzle 5', but the temperature of the hydrogen exhausted from the main nozzle 5 is high, so as a whole, (a) It is thought that there is no big difference from the case of .

FIG. 3 shows an embodiment in which part of the hydrogen from the pump 3 and the oxidizer (here oxygen) from the tank 7 are combusted in a gas generator 8, and the combustion gas is used to open the turbine 6'. show.

This cycle is almost the same as the cycle shown in FIG. 2(b). Although an oxidizing agent is required, it is possible to construct a cycle with greater flexibility than the cycle shown in FIG. 2(b).

The performance of the cooling and propulsion cycles of these examples will be specifically examined. Figure 4 shows the relationship between the flying Matsuha number and altitude used in this study. Matsuha 6 has an altitude of 20 km, Matsuha 8 has an altitude of 25 kI11, and Matsuha 10 has an altitude of 30 km.
m, and Matsuha 12 was assumed to be at an altitude of 35 km. This is close to the flight path at a dynamic pressure of 100 kPa.

The total length of the space plane is 70 m, the width is 10 m, the radius of curvature at the tip is 1 m, and the surface area requiring cooling is 600 m.
m2, the wall temperature on the air side is 10006K, the hydrogen temperature at the exit of the cooling channel is 800"K, and the emissivity of the aircraft surface is 0.5, then the coolant flow rate in the aircraft cooling cycle is It is shown in FIG.

The horizontal axis shows the flight Matsuha number, and the vertical axis shows the hydrogen flow rate. In the Matsuha 6, there is almost no heat exchange, but in the Matsuha 10, about 9 kg-5" of coolant is used. This is about 1/3 of the hydrogen flow rate of the air-breathing engine that generates 500 kN of thrust in the Matsuha 10. .

Figure 6 shows the pump outlet pressure. The results are obtained when the manifold pressure on the upstream side of the nozzle was set to Q, 5 MPa, and hydrogen necessary for cooling was pumped. The pressure difference in the turbine as shown in FIG. 2(a) is not included. The cooling flow path has a typical diameter of 2 square meters, and assumes a path that goes from the center of the fuselage to the tip and returns to the center of the fuselage.

In the Matsuha 6, the coolant flow rate is extremely small, so the pump outlet pressure is also low, but in the Matsuha 10, a pump outlet pressure of approximately 5 MPa is required.

When using the cycle shown in FIG. 2(a), for example, if the manifold pressure on the upstream side of the nozzle is set to 0.4 MPa and the turbine inlet pressure is set to 0.5 MPa, then 1.25 (=0
．． The cycle is sufficiently established at a turbine pressure ratio of 510.4). Since the pressure ratio is approximately 1, the temperature of the hydrogen exiting the turbine does not drop much to about 760'K.

When the cycle shown in FIG. 3 is used, the pump hydrogen flow rate will be greater than the present study result by the hydrogen flow rate required to drive the turbine. Therefore, the pump outlet pressure is expected to be slightly higher than the value shown in FIG.

Figure 7 shows the thrust when hydrogen is exhausted at 800' using a 10=1 nozzle. About 3 for hydrogen at 800'
A specific impulse of 70 seconds is obtained. Therefore, Matsuha 1o is about 3
A thrust of 0 kN is generated by this airframe cooling cycle. Using the cycle shown in Figure 2(a) as above,
The specific impulse at a turbine outlet temperature of 760' is approximately 360 seconds, and the thrust at Matsuha 10 is approximately 1 kN lower than at 800'.

As is well known in space planes, etc.

The air-breathing engine is cooled with liquid hydrogen, etc., but at high speeds, the amount of liquid hydrogen required for cooling exceeds the amount required for combustion, and if the entire amount is sent to the combustion chamber, the hydrogen/oxygen ratio collapses, causing engine performance to deteriorate. A tendency to decline occurs. Therefore, in many cases it is inappropriate to feed the hydrogen used for cooling the aircraft to the engine, and it is better to exhaust the heated hydrogen directly from the nozzle and use it as an auxiliary propulsion system. This auxiliary propulsion system is considered to be advantageous for use in controlling the attitude of the aircraft, reverse injection for braking, and the like.

(Effects of the Invention) This invention forcibly cools the aircraft using the on-board coolant as described above, which not only improves heat resistance against temperature increases due to aerodynamic heating, but also reduces the energy absorbed by the coolant. Since it can be actively used as an auxiliary propulsion system, highly efficient flight is possible.

[Brief explanation of the drawing]

Figure 1 is a conceptual diagram of the airframe cooling cycle when fuel is used as the coolant, Figure 2 is a conceptual diagram of the airframe cooling cycle when the pump drive system is a turbine, and Figure 3 is a conceptual diagram of the airframe cooling cycle when the pump drive system is a turbine. Figure 4 is a graph showing the relationship between flight Matsuha number and altitude, Figure 5 is a graph showing the coolant flow rate in the aircraft cooling cycle, and Figure 6
Figure 7 is a graph showing the pump outlet pressure of the cooling cycle.
The figure is a graph showing the thrust generated during the airframe cooling cycle. Symbols in the figure indicate the fuselage 2: Liquid hydrogen tank 3: Pump 4: Heat exchange section 5: Exhaust nozzle 5': Turbine exhaust nozzle 6: Pump drive device 6': Turbine 7:
Oxidizer tank 8: Shows gas generator. Patent Applicant Hideo Su, National Institute of Aerospace Science and Technology, Japan

Claims (1)

[Claims] 1) In a flying object that is exposed to high temperatures due to aerodynamic heating, etc., thrust is generated by cooling the airframe with a propellant and ejecting the heated coolant after this heat exchange from a nozzle. 2) An airframe cooling cycle according to claim 1, wherein the thrust obtained by the heated coolant is used for a propulsion system such as attitude control 3) The airframe cooling cycle according to claim 1, characterized in that the airframe cooling cycle is driven using heat obtained by airframe cooling.