JPS5950865B2 - Thermal propulsion device with super heater - Google Patents

Description

DETAILED DESCRIPTION OF THE INVENTION The present invention relates to a new and improved propulsion system, and more particularly to an electrothermal propulsion system for maneuvering space vehicles.

Electrically heated hydrazine propulsion devices are known and described in the following documents.

AIAA 8th Electric Propulsion Systems Conference, 1970, 70-
No. 1161.

AIAA/SAE 8th Propulsion Equipment Experts Meeting, 1972
Year, No. 72-1152.

AIAA 9th Electric Propulsion Systems Conference, 1972, 72-
No. 451 U.S. Pat. No. 3,081,595.

Gotsdad Space Flight Center Final Report NAS5-232
J. D. Kensley's paper published in 2002, “Study on single-wave propellants for electrothermal hydrazine propulsion systems J, March 1973-May 1974.

International Conference on the Properties of Hydrazine and Potential Uses of Hydrazine as an Energy Source (Poitiers, France, 21-25 October 1974)
Development of electrothermal hydrazine propulsion device for low thrust applications”
A paper entitled.

These propulsion devices have one or more of the following functions:

That is, the functions include attitude control, initial trajectory correction, initial position securing, attitude or trajectory correction, speed increase to compensate for deceleration due to drag, trajectory elevation, and obstacle avoidance maneuvering.

For example, commercial communications hygiene requires that all but the last three functions be accomplished by onboard propulsion systems.

Orbital weight 907.
In a 2 kg (2000 lb) spacecraft, for a 7 year mission, the onboard catalytic hydrazine system would weigh approximately 226.8 kg (500 lb).

The effective payload of observation equipment directly related to the purpose of the space vehicle is approximately 181.6 kg (approximately 400 pounds).
The remaining weight is allocated to power systems, structures, command equipment, ranging equipment, etc.

In addition to commercial communications satellites, military or scientific space vehicles are typical of long-life Earth-orbiting space vehicles that require maneuverability.

Any reduction in propulsion system weight increases payload, which is highly desirable.

Increasing the payload means increasing communication capabilities, increasing the accumulation of scientific data, and providing additional monitoring equipment, thereby making a significant difference to the invested capital of a spacecraft. It will give you a profit.

Initially, in the United States space program, virtually all spacecraft requiring onboard propulsion relied on cryogenic gas systems that produced a specific impulse of about 70 seconds.

However, as the missions of spacecraft become longer and more complex, the demands on spacecraft become more demanding.

To increase the required propulsion, the cold gases are supplemented by hydrogen peroxide, which is decomposed using a catalyst.

This increases complexity and presents various operational difficulties, but at the cost of significantly reducing the weight of the propulsion device.

However, hydrogen peroxide propulsion devices have been obsolete with the advent of catalytically decomposed anhydrous hydrazine, which further reduces weight and eliminates some of the problems in operating hydrogen peroxide.

This catalytic hydrazine propulsion device is typically 215-
Providing a specific impulse (steady state) of 235 seconds, such propulsion systems are currently in widespread use for space vehicles for a variety of missions.

Specific impulse is a commonly used performance indicator for the performance of a propulsion device, and is defined as the thrust per unit flow rate of propellant.

The specific impulse is determined by the energy content of the gas in the thrust chamber and the efficiency of the nozzle expansion process that converts the energy of this chamber into the kinetic energy of the exhaust gas.

The relationship between chamber energy and specific impulse Isp is Jsp-v’i It is.

Here, H is the portion of energy (enthalpy) that is converted into kinetic energy.

In a catalytic hydrazine propulsion system, anhydrous hydrazine is
It is thought that it decomposes into hydrogen and nitrogen in stages.

That is, the initial decomposition and the final decomposition as described below.

3N2H4 → 4NH3 + N2 + 144300 BTU (exothermic decomposition) 4NH3 → 2N2 + 6H2 - 79200 BTU (endothermic dissociation) These reactions result in approximately 870-980°C (1600-
A mixture of N2, N2, and NH3 is produced at a temperature of 1800°F.

Isentropic expansion of these decomposition products through the nozzle provides a specific impulse (theoretical value) of 210-260 seconds.

However, considering various performance loss mechanisms (e.g., heat loss, incomplete expansion, nozzle expansion, etc.), the steady-state specific impulse provided by current hydrazine propulsion systems is
A range of 15 to 235 seconds is typical.

Because the chamber enthalpy (and theoretical thrust) is determined by the net chemical energy released by propellant decomposition, the potential for performance improvement in catalytic hydrazine propulsion devices is extremely limited.

There is therefore a need to increase the propulsion capacity to at least 300 Isp while maintaining a more or less constant power supply without significantly increasing the weight of the propulsion unit or propellant.

The method and apparatus of the present invention can meet this need, and an electrothermal propulsion device implementing the method of the present invention includes an injection tube, a chamber to be heated, and a screen pack assembly ( a superheater, a superheater heat source, and an exhaust nozzle.

The heater chamber operates in thermal, catalytic, and quasi-catalytic modes.

In the thermal mode, the heater may be an electrically actuated coil helically wrapped around the chamber, a radioisotope source, or the like.

A screen pack temperature of 200° C.-450° C. is required when operating in catalytic mode; above 450° C., the propulsion device is operated in pyrolysis mode.

This requires high power but improves propulsion device life and specific impulse.

The screen pack (catalyst bed) provides uniform decomposition.

A typical screen pack is made of platinum wire.

When using hydrazine with an inlet pressure of 1.724 MN/M2, the diameter is 0.5 cm, 52 meshes, and the wire diameter is 0.5 cm.
60 screens of 1 mm are suitable.

Moreover, it may be replaced with a composite type screen.

A retaining screen with a wire diameter of 0.28 mm can be used with 40 mesh.

A superheater continuously and stably provides energy to the gas released from the heater chamber, and provides a continuous, steady, and approximately constant output.

The superheater increases the temperature of the gas in the chamber.

As a vortex superheater, the tangential propellant injection system creates a steep radial temperature gradient, creating a relatively cool gas layer near the walls of the cavity.

Therefore, conventional insulation can be used if necessary.

Propulsion devices use single-wave propellants (liquid propellants that decompose thermally or catalytically into a gaseous state; typical propellants include anhydrous hydrazine, ammonia, water, hydrazine azide, monomethylhydrazine, and asymmetric dimethylhydrazine. , mixtures thereof, etc., but hydrazine is the preferred fuel.

The net overall effect is that the exit temperature of the propellant through the nozzle is very close to that of the vortex heating element, and the weight of the propulsion system can be increased by only a few ounces to increase thrust capacity by 2.
Increase from 00-235 Isp to approximately 300-340 Isp.

This corresponds to a reduction in propellant requirements of approximately 20-25%.

For example, for a 907.2 kg (2000 lb) spacecraft payload that typically requires 226.8 kg (500 lb) of propellant, the propulsion system of the present invention reduces this propellant requirement to 170-188 kg. (about 37
5i15 pounds).

The invention will be easily understood from the drawings.

FIG. 1 shows a propulsion device having a cylindrical heater chamber 11. FIG.

Heat is supplied to the heater chamber and the nozzle by a helical electrical element 12 threaded onto the outside of the heater chamber.

A platinum screen pack is placed within the heater chamber, which absorbs heat from the electrical elements and causes fuel, such as hydrazine, to combust when it comes into contact with the screen pack as it passes through the chamber.

Fuel is supplied from the valve to the heater chamber through the injection pipe 15.

This tube 15 extends through and is in contact with a perforated thermal barrier tube 16.

If desired, the valve may be heated by any suitable means.

The super heater 22 shown in FIG. 1 is arranged at 90° with respect to the heater chamber 11.

The superheater 22 has a cylindrical chamber wall 23 , which is arranged at an angle of 90° to the longitudinal axis of the heater chamber 11 .

A coiled heater 24 is placed along the axis of the chamber and is heated by a power source 25.

If desired, decomposition and/or heating can be performed by arc plasma or electrical discharge.

At the outlet of the superheater a nozzle 26 is provided, which may be formed as a separate element or as an integral part of the chamber wall 23.

If desired, the entire propulsion device and superheater can be housed within the metal casing 27.

Decomposed propellant gas from the heater chamber is fed tangentially into the superheater section and further heated to provide additional propulsion capacity.

The tangential injection of the decomposed propellant gas into the superheater creates a vortex of gas between the heater and the chamber wall around the central heater within the superheater.

This results in the following important advantages:

First, the swirling propellant gas flow through the superheater increases the flow path length, and the flow path length within the superheater increases compared to the flow path length and residence time when the gas flows longitudinally past the heater. and substantially increase residence time.

Therefore, in a vortex flow, the time during which the propellant gas is subjected to superheating, and thus the temperature rise of the gas created by superheating, is substantially increased.

Second, as a vortex of propellant gas is created between the center heater and the superheater chamber wall, the gas picks up thermal energy that radiates from the center heater toward the chamber wall.

This maximizes the thermal energy transferred to the gas and the gas temperature, and minimizes heat loss through the chamber walls.

The tangential injection of the decomposed propellant gas into the superheater and the placement of the heater in the center of the chamber together reduce heat loss and increase the exit temperature of the propellant gas, thereby increasing the Increase specific impulse.

FIG. 2 shows a spacecraft 30 that uses a propulsion system to perform various positioning maneuvers.

In the illustrated embodiment, a twin propulsion device 3 according to the invention
1.32 are mounted on opposite sides of the spacecraft for north-south maneuvers.

Typically, a minimum room temperature of about 800°F (430°C) is required for propulsion device operation.

Warm-up time is 1 with a heater input of 5-8 watts.
It is less than 0 minutes.

The propellant valve is opened by ground command. heat exchanger 2
4, the propulsion device temperature equilibrates in about 25 seconds and the specific impulse is 200-205 seconds.

The heat exchanger output can be activated by ground command or can be switched on simultaneously with the control valve.

At full flow rate, the thermal response of the vortex superheater is several seconds, with a corresponding transient specific impulse.

The thermal response of a superheater is 100% longer.

In either configuration, the instructions can be interlocked to prevent application of heat exchanger power without propellant flow.

The preferred temperature level at the outlet of the heater chamber is about 150
0°F (815°C) to 2000°F (1093°C)
), and the preferred nozzle temperature of the superheater is approximately 250°C.
0°F (1370°C) or higher.

The preferred nozzle temperature range is 2500°F (1370°C
) to 5000°F (2760°C), above which the upper temperature limit is limited only by the selected materials of construction.

In the graph in Figure 3, the vertical axis on the left is the specific impulse generated by the propulsion device in vacuum (outer space), and the thrust generated by the propulsion device is 0.0.
The power input to the heater of the propulsion device, measured in watts per kg, is plotted on the right vertical axis, and the gas temperature measured at the nozzle outlet is plotted on the horizontal axis.

The arrow points to the vertical axis along which the curve of each graph is read.

The graph of FIG. 3 shows that the actual specific impulse and the theoretically available specific impulse from hydrazine match very closely.

Furthermore, from this graph, there is no irregularity at any point,
It can be seen that the specific impulse is constant.

The propulsion device of the present invention has a payload of 907.2 kg (200
The amount required to produce a specific impulse of 215-235 seconds using a device according to this technology, as well as to provide a specific impulse of 300-340 seconds, sufficient for fixing the north-heavy position in an artificial sanitary vessel of 0 lbs. Approximately 38.5kg to 56.7k compared to
g (approximately 85 to 125 pounds) of propellant can be saved.

Since heaters and superheaters only weigh a few ounces,
The extra weight of the fittings is negligible.

This reduced propellant weight advantage is significant for modern space vehicles.

[Brief explanation of drawings]

FIG. 1 is a partially cutaway perspective view showing an embodiment of the present invention. FIG. 2 is a perspective view of the sanitary body showing the structure using dual sets of propulsion devices for north-south maneuvering. FIG. 3 is a graph showing changes in specific impulse, gas temperature at the nozzle outlet, and electric power to the heater. 11... Heater chamber, 12... Spiral electric element, 22... Super heater, 23...
... Wall of cylindrical chamber, 24 ... Coiled heater, 2
5...Power supply, 26...Nozzle, 27.
...Metal casing, 30... Space vehicle, 31, 32... Propulsion device.

Claims (1)

[Scope of Claims] 1- Supplying a liquid propellant to a decomposition chamber; Decomposing the single-wave propellant in the decomposition chamber into gaseous decomposition products at a decomposition temperature; The high temperature gaseous decomposition products are injected into the superheater chamber between the heater and the wall of the superheater chamber, around the heater and along it into a swirling flow through the superheater chamber and away from the heater during an extended stagnation period within the superheater chamber. The radiant heat to the wall is absorbed by the decomposition products and the gaseous decomposition products are released from the superheater chamber through the nozzle of the superheater chamber to create a thrust; and the said decomposition products pass through the nozzle to a temperature above the decomposition temperature. A method for generating thrust, characterized in that the heater is operated at a temperature such that the heater is emitted at a temperature of . 2. a decomposition chamber having an inlet for receiving the single-wave propellant to be decomposed; means within the decomposition chamber for decomposing the single-wave propellant into gaseous decomposition products at an initial decomposition temperature; a nozzle at one end on the shaft; a superheater chamber having a wall around an axis; a heater disposed within the superheater chamber on said axis radially spaced relative to the wall of the superheater chamber; a high temperature from said decomposition chamber to said superheater chamber; The heater includes means for directing cracked gas and injecting the high temperature cracked products tangentially into the superheater chamber, the heater being configured to discharge the cracked products through the nozzle at a temperature above the cracking temperature. A thrust generating device characterized by being able to operate at temperatures such as