JPH052559B2 - - Google Patents

Description

Detailed Description of the Invention (Field of Industrial Application) This invention relates to the cooling cycle of a high-speed flying vehicle in which the fuselage is exposed to high temperatures, and in particular to the cooling cycle of a high-speed flying vehicle in which the fuselage is exposed to high temperatures. Regarding the propulsion device by.

(Prior technology) When a space shuttle re-enters the atmosphere, the aircraft is exposed to high temperatures of up to 2,000 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.

(Problems to be Solved by this Invention) In the case of conventional space shuttle re-entry, the time of exposure to high temperatures is not that long, and the heat insulating layer prevents the aircraft from being directly exposed to high temperatures. However, in the case of supersonic intercontinental vehicles, etc., the time when they are subjected to aerodynamic heating is long, and it becomes difficult to sufficiently prevent the temperature rise of the aircraft with just a heat insulating layer. Furthermore, the protection of the fuselage by the heat insulating layer merely blocked the generated heat and allowed it to escape, but did not actively utilize it.

(Means for Solving the Problems) In this invention, in high-speed flying objects such as space planes, the airframe is exposed to high temperature air due to aerodynamic heating, etc. In order to protect the airframe from this high temperature, It is cooled by a propellant such as liquid 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.
It is preferable to use the coolant used for cooling the aircraft to drive 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 a drive device for the pump 3, but any drive device may be used for this drive.

FIG. 2 shows an embodiment in which the pump drive system is a turbine. Figure 2a shows the case where the entire amount of hydrogen after heat exchange drives the turbine 6' and then exhausts the hydrogen through the nozzle, and Figure 2b shows the case where part of the hydrogen after heat exchange is used to drive the turbine 6'. , the remainder 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 with 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.
Overall, it is considered that there is no major difference from case a.

FIG. 3 shows a gas generator 8 combusting part of the hydrogen from the pump 3 and an oxidizing agent (oxygen here) from the tank 7, and the combustion gas is used to feed the turbine 6'.
An example will be shown in which the drive is performed.

This cycle is approximately the same as the cycle shown in Figure 2b. Although an oxidizing agent is required, a cycle can be constructed that is more flexible than the cycle shown in Figure 2b.

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, assuming an altitude of 20 km for Matsuha 6, 25 km for Matsuha 8, 30 km for Matsuha 10, and 35 km for Matsuha 12.
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, the surface area that requires cooling is 600 ^m2 , the air side wall temperature is 1000°K, and the hydrogen temperature at the exit of the cooling channel is 800°K. Figure 5 shows the coolant flow rate in the airframe cooling cycle when the temperature is 0.5°K and the emissivity of the airframe surface is 0.5.

The horizontal axis shows the flight number and the vertical axis shows the hydrogen flow rate. Matsuha 6 hardly exchanges heat,
Matsuha 10 uses approximately 9 kg・s ^-1 of coolant. This is approximately 1/3 of the hydrogen flow rate of the Matsuha 10's air-breathing engine, which generates a thrust of 500 kN.

Figure 6 shows the pump outlet pressure. This is the result when the manifold pressure on the upstream side of the nozzle was set to 0.5 MPa and the hydrogen necessary for cooling was pumped. Figure 2a
It does not include the pressure difference at the turbine as shown in . The cooling flow path has a typical diameter of 2 cm, and assumes a path from the center of the fuselage to the tip and back 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, approximately
A pump outlet pressure of 5MPa is required.

When using the cycle shown in Figure 2a, for example, if the nozzle upstream manifold pressure is 0.4 MPa and the turbine inlet pressure is 0.5 MPa, then 1.25 (=
The cycle is sufficiently established at a turbine pressure ratio of 0.5/0.4). Since the pressure ratio is approximately 1, the temperature of the hydrogen leaving the turbine does not drop much, at approximately 76°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 at 800°K is exhausted using a 10:1 nozzle. Hydrogen at 800°K gives a specific impulse of about 370 seconds. Therefore, in the Matsuha 10, approximately 30kN of thrust is generated by this airframe cooling cycle. When the cycle shown in Figure 2a is used as described above, the specific impulse at a turbine outlet temperature of 760°K is approximately 360 seconds, and the thrust at Matsuha 10 is approximately 1 kN lower than at 800°K.

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, so the entire amount is pumped into the combustion chamber. This tends to cause a decline in engine performance due to the collapse of the hydrogen/oxygen ratio. 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 effective for use in aircraft attitude control, reverse injection for braking, etc.

(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 drawings]

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,
Figure 3 is a conceptual diagram of the airframe cooling cycle when using a gas generator, Figure 4 is a graph showing the relationship between the flight Matsuha number and altitude, and Figure 5 is a graph showing the coolant flow rate in the airframe cooling cycle. Figure 6 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. The symbols in the diagram are: 1...Airframe, 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
...Indicates a gas generator.

Claims (1)

[Claims] 1. In a flying object that is exposed to high temperatures due to heat applied to the airframe during flight, the airframe is cooled by a propellant, and the heated coolant after this heat exchange is performed. A propulsion device that uses body cooling heat to obtain thrust by ejecting from a nozzle. 2. A propulsion device using body cooling heat according to claim 1, wherein the thrust obtained by the heated coolant is used for a propulsion system such as attitude control. 3. A propulsion device using body cooling heat according to claim 1 or 2, characterized in that the heat obtained by cooling the body is used to drive the propulsion device using body cooling heat.