Power plants – Reaction motor – Method of operation
Reexamination Certificate
2001-01-02
2001-10-30
Freay, Charles G. (Department: 3746)
Power plants
Reaction motor
Method of operation
Reexamination Certificate
active
06308514
ABSTRACT:
The present invention refers to a rocket engine nozzle with an outlet portion or thrust chamber having a curved profile in axial section.
During start-up and stop transients in sea-level rocket engines significant dynamic Loads usually occur. These loads are generally attributed to the disordered flow characteristics of the flow during flow separation.
The outlet portion of nozzles for liquid propellant rocket engines often operate at conditions where the main jet exhausts into a non-negligible ambient pressure. Examples of such rocket engines are large liquid propellant sea-level rocket engines for boosters and core stages that are ignited at sea-level and upper stage rocket engines ignited during stage separation.
Dynamic loads are due to the instationary nature of the thrust chamber flow during star and stop transients and during steady-state operation with separated flow in the nozzle. The rules for and effects of such flow separation have been studied and presented at the 30th AIAA/ASME/SAE/ASEE Joint Propulsion Conference, June 1994, Indianapolis, Ind., USA in a paper “Aero-elastic Analysis of Side Load in Supersonic Nozzles with Separated Flow”, Volvo Aero Corporation, to which may be referred. The dynamic loads are generally of such a magnitude, e.g. of the order of 50-100 kN, that they present life-limiting constraints for the design of thrust chamber components. These constraints result in higher weight for thrust chamber structural elements. Furthermore the largest possible area ratio that can be used on the nozzle extension is limited by the requirement of attached flow during steady-state operation.
The final consequences of the dynamic loads are constraints on the overall performance-to-weight ratio of the thrust chambers and a subsequent limitation of the amount of cay-load that can be delivered into orbit by the rocket launcher.
For eliminating the drawbacks of prior nozzles a great number of techniques have been suggested which all, however, have turned out to have themselves significant drawbacks in various respects. The main difficulties refer to the function, performance, cooling and reliability.
Thus traditional bellshaped nozzles give a limited function and substantial start and stop transient loads. A dual bell nozzle also suffers from severe transient dynamic loads. External expansion nozzles have been suggested but not been sufficiently tested. A bell nozzle equipped with trip rings reduces dynamic loads but with too large a performance loss. Said nozzles also suffer from difficult cooling problems. Finally, extendible nozzles and ventilated nozzles have been suggested but both require mechanisms with functions that are not possible to verify prior to flight.
The main object of the present invention now is to suggest a rocket engine nozzle structure which provides for an advantageous flow control within the diverging outlet portion of the nozzle which makes it possible to reduce the weight of rocket engine nozzle and to gain increased performance.
According to the invention this is achieved by a nozzle which is substantially distinguished in that, circumferentially, said outlet portion in axial section has a radius, the length of which varies. Preferably, the radius length varies periodically and most preferably periodically so as to create a polygonal circumferential shape of the nozzle.
According to the invention, the separated flow can be controlled satisfactorily by this very limited non-axisymmetric modification of the nozzle wall contour. This modification exhibits no significant negative effects on performance, reliability, cooling and manufacturing aspects of the nozzle outlet portion.
The invention thus provides for the design a sea-level rocket engine nozzle with significantly higher vacuum performance through a larger nozzle area ratio and reduced weight.
REFERENCES:
patent: 3292865 (1966-12-01), Short et al.
patent: 3305177 (1967-02-01), Fage
patent: 4707899 (1987-11-01), Singer
patent: 5343698 (1994-09-01), Porter et al.
Häggander Jan
Pekkari Lars-Olof
Freay Charles G.
Ostrolenk Faber Gerb & Soffen, LLP
Torrente David J.
Volvo Aero Corporation
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