Thrust chamber assembly

Details Thrust chamber assembly Thrust chamber assembly

2002-04-01

2004-03-02

Gartenberg, Ehud (Department: 3746)

Power plants

Reaction motor

Method of operation

C060S257000

Reexamination Certificate

active

06698184

ABSTRACT:

BACKGROUND AND SUMMARY OF THE INVENTION
The invention concerns a thrust-chamber assembly for space rocket engines that are used particularly for adjusting the position of a satellite.
Space rocket engines have a fuel injector head through which fuel and oxidizer are injected into the thrust chamber. During operation, a film of fuel and oxidizer with a constant thickness of about 0.5 mm builds up on the inner wall of the precombustion chamber where these materials are mixed and prevents overheating of the precombustion-chamber wall. In practice, this film has proved unstable at times, causing the temperature of some areas of the precombustion-chamber wall to rise and reduce its stability.
This film, which consists of combustible materials, i.e. fuel and oxidizers, that have not yet ignited, proved unstable particularly during pulsed operation, when a considerably smaller amount of combustible material is injected into the combustion chamber than during stationary operation, as this marginal cooling film is caused to evaporate by the heat stream from either the narrowest diameter of the nozzle, which is considerably hotter, or the narrow part of the nozzle.
Some of the current documentation on this technology discusses means of dealing with high temperatures in rocket engines.
U.S. Pat. No. 3,719,046 describes how the combustion chamber of a rocket engine is surrounded by a heat-conductor containing a wick-like material saturated with a fluid, this material being located adjacent to the combustion chamber. The fluid evaporates when it absorbs heat from the combustion chamber, and the vapor is delivered to a heat exchanger, where it condensed. The fluid is then returned to the wick-like material.
WO 96/25596 describes a combustion chamber for a rocket engine with a lining of rhodium, iridium, or a rhodium-iridium alloy that is resistant to high temperatures, and also guarantees resistance to corrosion from unburned fuel elements.
However, these documents offer no solution to the problems described above, which are connected to the cooling film in space rocket engines.
The purpose of this invention is therefore to shape the thrust chamber in such a way that the combustible materials fed into the combustion chamber of a satellite engine via the injector head will be certain to form a stable and homogenous film on the inner wall of the combustion chamber in the proximity of the injector head.
The purpose of this invention is therefore to configure the combustion chamber in such a way that the combustible materials fed into the combustion chamber of a satellite engine via the injector head will be certain to form a stable and homogenous film on the inner wall of the combustion chamber in the proximity of the injector head.
This problem is dealt with using the features described in claim
1
. Alternative methods are described in the sub-claims.
One advantage of the solution used by this invention is that existing forms of thrust-chamber construction remain essentially unchanged. Better operational reliability can be achieved through simple modifications.


REFERENCES:
patent: 3719046 (1973-03-01), Sutherland
patent: 5780157 (1998-07-01), Tuffias et al.
patent: 6151887 (2000-11-01), Haidn et al.
patent: 4137638 (1993-06-01), None
patent: 0780564 (1997-06-01), None
patent: 96/25595 (1996-08-01), None
Sutton, G.P. and Ross, D.M., “Rocket Propulsion Elements”, John Wiley and Sons, New York, 1976, pp. 278, 297.*
Copy of the International Search Report.

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