Patent Publication Number: US-6655127-B2

Title: Pre-burner for low temperature turbine applications

Description:
This application is a is a divisional of application of application Ser. No. 09/761,957 filed on Jan. 17, 2001 now U.S. Pat. No. 6,505,463. 
    
    
     BACKGROUND OF THE INVENTION 
     1. Field of the Invention 
     This invention relates generally to turbopump fed rocket engines, and, more particularly, to a fuel rich pre-burner that operates at a very low mixture ratio to drive a turbine in a hydrogen peroxide based rocket engine. 
     2. Discussion of the Related Art 
     Rocket engines typically employ liquid propellants which are burned to form hot gases. The high pressure hot gases are then expanded though a specially shaped nozzle, thereby producing a thrust force for the rocket. The propellants usually consist of a liquid fuel and a liquid oxidizer. In at least one well known hydrogen peroxide based engine design, the liquid oxidizer may also be used to drive a turbine. In this case, a portion of the liquid oxidizer is decomposed into an oxidizing gas by passing it though an appropriate catalyst. The oxidizing gas is then used to drive the turbine which in turn may be used to drive a fuel pump and/or a oxidizer pump. 
     Hydrogen peroxide is at least one liquid oxidizer commonly used in turbo pump fed rocket engines. However, as the demands for rocket engine performance increase, higher concentrations of hydrogen peroxide are needed to meet these demands. Unfortunately, high concentrations of hydrogen peroxide (e.g., &gt;92%) produce gas temperatures that exceed the temperature and oxidation limits of traditional materials used for turbines. In order to use high concentrations of hydrogen peroxide, designers will need to develop new materials which can withstand the higher temperature gases that are passed through the turbine. 
     Therefore, it is desirable to provide a turbopump fed rocket engine that can utilize high concentrations of hydrogen peroxide or other monopropellants that decompose at high temperatures. By utilizing a fuel rich pre-burner that operates at a very low mixture ratio to drive the turbine, the temperature of the drive gas can be maintained at moderate levels, thereby enabling the use of conventional materials for the turbine. Thus, high performance turbopump fed rocket engines can be developed independent of the creation of a new high temperature resistant turbine material. 
     SUMMARY OF THE INVENTION 
     In accordance with the teachings of the present invention, a method is provided for using monopropellant oxidizers having high decomposition temperatures to drive a turbine in a turbopump fed rocket engine. The method includes the steps of: (a) providing rocket fuel to a preburner; (b) providing a portion of the oxidizer to the preburner; (c) converting the fuel and the oxidizer into a fuel rich gas; and (d) passing the fuel rich gas through a turbine, thereby using at least a portion of the oxidizer to drive the turbine in a rocket engine. 
    
    
     Additional objects, features and advantages of the present invention will become apparent from the following description and appended claims taken in conjunction with the accompanying drawings. 
     BRIEF DESCRIPTION OF THE DRAWINGS 
     FIG. 1 is a block diagram of a prior art propellant supply system for a turbopump fed rocket engine; and 
     FIG. 2 is a block diagram of a propellant supply system for a turbopump fed rocket engine in accordance with the present invention. 
    
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS 
     FIG. 1 illustrates a propellant supply system  10  for a conventional turbopump fed rocket engine  12 . The propellant supply system  10  generally includes a fuel pump  14 , an oxidizer pump  16 , a gas generator  18  and a turbine  20 . The engine  12  is further defined to include an injector  22 , and a main combustion chamber  26 . 
     Liquid rocket fuel is typically stored at low pressures in a fuel tank (not shown). The fuel pump  14  provides pressurized fuel from the fuel tank to the injector  22  of the engine  12 . Prior to reaching the engine, the pressurized fuel may pass through at least one orifice  32  and a main fuel throttle valve  34 . 
     Likewise, liquid oxidizer is stored at low pressures in an oxidizer tank (not shown). The oxidizer pump  16  provides pressurized liquid oxidizer to gas generator  18 . Prior to reaching the gas generator  18 , the liquid oxidizer may pass through at least one orifice  42 , a main oxidizer throttle valve  44  and through a cooling jacket  46  that is formed along the outer surface of the combustion chamber. In this way, the liquid oxidizer may be used to cool the high temperatures associated with the main combustion chamber  26  of the engine  12 . 
     The liquid oxidizer is then decomposed into an oxidizing gas by passing it through an appropriate catalyst inside the gas generator  18 . The oxidizing gas is used to drive the turbine  20 , which in turn is used to drive the fuel pump  14  and/or the oxidizer pump  16 . Lastly, the pressurized fuel and the oxidizing gases are injected by the injector  22  into the main combustion chamber  26  of the engine  12 . 
     Hydrogen peroxide is at least one liquid oxidizer commonly used in the above-described turbo fed rocket engine. In operation, moderate concentration levels of hydrogen peroxide (i.e., &lt;92%) produce gas temperatures within the temperature and oxidation limits of the conventional materials used for the turbine  20 . In contrast, higher concentration levels of hydrogen peroxide produce gas temperatures that exceed the temperature and oxidation limits of the materials used for the turbine  20 . For example, a hydrogen peroxide liquid having a concentration level of 98% will typically result in an oxidized gas having a temperature in the range of 2200 degrees Rankin as it passes through the turbine  20 . 
     In accordance with the present invention, a propellant supply system is provided that can utilize high concentrations of hydrogen peroxide as the liquid oxidizer. As will be more fully described below, the propellant supply system incorporates a fuel rich pre-burner that operates at a very low mixture ratio in order to drive a turbine. In this way, the temperature of the drive gas can be maintained at moderate levels, thereby enabling the use of conventional materials for the turbine. 
     A first preferred embodiment of the fuel supply system  10 ′ is shown in FIG.  2 . While the following description is provided with reference to hydrogen peroxide as the oxidizer, it is readily understood that the broader aspects of the present invention are also applicable to other monopropellant oxidizers that decompose at high temperatures. Exemplary monopropellant oxidizers may include (but are not limited to) ethylene-oxide, nitromethane and HAN. 
     As previously described, the fuel supply system  10 ′ includes a fuel pump  14 ′, an oxidizer pump  16 ′, and a turbine  20 ′. However, in this embodiment, the fuel pump  14 ′ provides pressurized fuel to a first gas generator  50  (alternatively, referred to as a fuel rich pre-burner). In order to reach the gas generator  50 , the pressurized fuel may pass through at least one orifice  32 ′ and a main fuel throttle valve  34 ′. For illustration purposes, kerosene is used as the rocket fuel (e.g., JP-8). However, it is envisioned that other types of well known rocket fuels (e.g., propane, hydrazine, methane, etc.) are also within the scope of the present invention. 
     In addition, a relatively small portion of the pressurized hydrogen peroxide is directed to the first gas generator  50 . For instance, if the flow rate of hydrogen peroxide at the outlet of the oxidizer pump  16 ′ is 33.9 lb/s, then a reduced flow rate of 4.38 lb/s may be directed to the first gas generator  50 . As will be apparent to one skilled in the art, the precise flow rate of the oxidizer to the first gas generator  50  is dependent on the nature of the two propellants as well as other system parameters. 
     The first gas generator  50  operates at a very low mixture ratio to convert (i.e., burn) the liquid propellants into a fuel rich gas. In contrast to the conventional approach, the temperature of the fuel rich gas can be maintained at moderate levels. For instance, a gas generator receiving fuel at a flow rate of 4.86 lb/s at 1880 psi and hydrogen peroxide at a flow rate of 4.38 lb/s at 1850 psi may output a fuel rich gas having a temperature in the range of 1710 degrees Rankin. At this moderate temperature level, the fuel rich gas may be used to drive the turbine  20 ′. The turbine  20 ′ is in turn used to drive the fuel pump  14 ′ and/or the oxidizer pump  16 ′. After passing through the turbine  20 ′, the fuel rich gas is directed to the injector  22 ′ associated with the engine  12 ′. 
     The remaining portion of the hydrogen peroxide is directed to the engine  12 ′. In this circuit, the hydrogen peroxide may pass through at least one orifice  42 ′ and a main oxidizer throttle valve  44 ′. At the engine, the hydrogen peroxide passes through a cooling jacket  46 ′ that is formed along the outer surface of the combustion chamber. In this preferred embodiment, the hydrogen peroxide may enter the cooling jacket  46 ′ having a temperature in the range of 530 degree Rankin, but exits the cooling jacket  46 ′ having a temperature in the range of 630 degree Rankin. Thus, the hydrogen peroxide may be used to cool the main combustion chamber  26 ′ of the engine  12 ′. 
     The second gas generator  24 ′ is then used to decompose the heated hydrogen peroxide into an oxidizing gas before it is injected into the combustion chamber  26 ′. Lastly, the fuel rich gas and the oxidized gases from the second gas generator  24 ′ are injected by the injector  22 ′ into the main combustion chamber  26 ′ of the engine  12 ′. 
     The foregoing discussion discloses and describes merely exemplary embodiments of the present invention. One skilled in the art will readily recognize from such discussion, and from the accompanying drawings and claims, that various changes, modifications and variations can be made therein without departing from the spirit and scope of the invention as defined in the following claims.