Patent Abstract:
An engine system includes a thrust chamber that has a cooling channel. The cooling channel is adapted to provide sustained cracking conditions for a fluid at steady-state operating conditions. A turbine has an input in fluid communication with an output of the cooling channel. A pump is mechanically coupled with the turbine and is in fluid communication with the cooling channel.

Full Description:
BACKGROUND 
     This disclosure relates to a thrust chamber of a rocket engine system that allows higher energy from hydrocarbon fuels. 
     Bi-propellant rocket engines are known and used to power aerospace vehicles. A typical bi-propellant rocket engine can utilize an expander cycle. The expander cycle typically involves heating the fuel, which is then expanded over a turbine drive system to drive a propellant pump before delivery to the combustion chamber. 
     Typically, the expander cycle fuel is a light-molecule fuel, such as liquid hydrogen, methane or propane. The expander cycle fuel has a high specific heat that is advantageous to cooling the chamber and/or nozzle and providing the energy to power the propellant pumps. Heavier molecule hydrocarbon fuels have not found widespread use in expander cycle rocket engines because at high temperatures, heavier fuels tend to form coke deposits that block the passages and foul the system. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The various features and advantages of the disclosed examples will become apparent to those skilled in the art from the following detailed description. The drawings that accompany the detailed description can be briefly described as follows. 
         FIG. 1  illustrates an example aerospace engine system. 
         FIG. 2  illustrates an example thrust chamber. 
     
    
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT 
       FIG. 1  schematically illustrates selected portions of a rocket engine system  22 . As will be described, the rocket engine system  22  is designed to allow the use of hydrocarbon fuels. Although depicted with a particular geometry and arrangement, it is to be understood that the concepts described herein are not limited to use with the specific rocket engine system  22 . 
     The illustrated rocket engine system  22  includes a thrust chamber  24  having walls  26  that define a combustion section  28 , a throat section  30  and a nozzle section  32 . In general, the combustion section  28 , the throat section  30  and the nozzle section  32  form an hourglass shape. That is, the combustion section  28  is relatively wide and narrows to the throat section  30 , which then widens to the nozzle section  32 . As shown, the nozzle section  32  is bell shaped. 
     The walls  26  of the thrust chamber  24  include cooling passages  34  therein. As shown in the illustration of the thrust chamber  24  in  FIG. 2 , the walls  26  are constructed from tubes or passages arranged side-by-side to form the hourglass shape of the thrust chamber  24 . The interiors of the tubes or passages serve as the cooling passages  34  through which fuel flows to cool the thrust chamber  24 . 
     A fuel pump  36  in the rocket engine system  22  delivers fuel to the thrust chamber  24 . In that regard, a fuel passage  38  fluidly connects the thrust chamber  24  and the fuel pump  36 . The fuel passage  38  splits into sub-passages, with a first sub-passage  38   a  leading to the combustion section  28  of the thrust chamber  24  and bypassing the cooling passage  34 . A second sub-passage  38   b  leads to the cooling passage  34  of the thrust chamber  24 . 
     In embodiments, the second sub-passage  38   b  continues on from the cooling passage  34  to a turbine  40 , which is coupled to drive the fuel pump  36 . From the turbine  40 , the second sub-passage  38   b  leads to the combustion section  28  of the thrust chamber  24 . Alternatively, the fuel from the turbine  40  may be dumped overboard instead of going to the combustion section  28 . An additional pump  42  may also be coupled with the turbine  40  to deliver oxidizer to the combustion section  28  through an oxidizer passage  44 . 
     In embodiments, the cooling passage  34  may include a catalytic material  48  that chemically interacts with fuel flowing through the cooling passage  34 . The catalytic material  48  may be a catalytic coating that lines the interior walls of the cooling passage  34 . The catalytic coating composition and/or conditions within the cooling passages (pressure, temperature, etc) are established to provide an environment sufficient to sustain cracking of the hydrocarbon fuel selected. The condition and catalyst will vary depending on the hydrocarbon selected as well as pertinent engine and thrust chamber characteristics. 
     The arrangement of the rocket engine system  22  and thrust chamber  24  allows the use of hydrocarbon fuels, such as kerosene. As an example, kerosene can form coke deposits at the temperatures (approximately 1300.degree. F./704.degree. C. or greater) experienced in the cooling passages  34  of a conventional thrust chamber. However, controlling the fuel flow rate, pressure and/or temperature with the use of the catalytic material (not shown), a reduction of coking can be achieved. The reduced coking allows such fuels to be used as a propellant in the rocket engine system  22  without coke deposits that could otherwise block the fuel passages and foul the turbine. The cracking process itself is endothermic, and thereby improves the cooling capability of the hydrocarbon fuel to the advantage of the engine cycle, for example, enhanced cooling, energizing the fuel delivered to the turbine(s), and increasing the energy content of the fuel delivered to the thrust chamber. 
     In embodiments, the fuel is initially a liquid that is delivered through the fuel passage  38  from the fuel pump  36 . The split in the fuel passage  38  diverts a portion of the liquid fuel through the first sub-passage  38   a  and another portion of the liquid fuel through the second sub-passage  38   b . The ratio of the flow split is determined to provide sufficient fuel to cool the thrust chamber while sustaining the conditions required for cracking the hydrocarbon fuel in the cooling passages  34 . 
     Optionally, a flow splitter  50  is provided within the fuel passage  38  to control the split of flow of the fuel. In that regard, a controller  52  in communication with the flow splitter  50  may command the flow splitter  50  to control the ratio of flow to each sub-passage  38   a ,  38   b . The controller  52  may also be in communication with the other control valves as desired to control rocket engine system  22 . 
     With the split in the fuel passage  38 , only a portion of the fuel flows through the cooling passage  34 , while the other portion flows directly to the combustion section  28 . By controlling the amount of fuel that flows through the cooling passage  34 , the controller  52  can ensure that the fuel in the cooling passage  34  heats to a predetermined temperature to sustain steady-state cracking of the hydrocarbon fuel in the cooling passages  34  prior to injection into the turbine  40 . That is, by reducing the amount of fuel delivered to the cooling passage  34 , the fuel flowing through the cooling passages  34  can be sustained above a critical temperature in a steady state operating condition for cracking and subsequent expansion in the turbine  40  to drive the fuel pump  36 . 
     Additionally, the catalytic material  48  within the cooling passage  34  serves to crack the heated fuel into lighter molecules thereby reduce coking of the fuel. Furthermore, the chemical cracking of the fuel is an endothermic reaction that absorbs additional heat from the thrust chamber  24 . Also, the conversion of the fuel into lighter molecules facilitates converting the fuel into a gaseous state for expansion over the turbine  40 . The rocket engine system  22  thereby allows the use of relatively heavy hydrocarbon fuels, such as kerosene. The fuel thereby serves the dual purposes of cooling the thrust chamber  24  and driving the turbine  40 . 
     Although a combination of features is shown in the illustrated examples, not all of them need to be combined to realize the benefits of various embodiments of this disclosure. In other words, a system designed according to an embodiment of this disclosure will not necessarily include all of the features shown in any one of the Figures or all of the portions schematically shown in the Figures. Moreover, selected features of one example embodiment may be combined with selected features of other example embodiments. 
     The preceding description is exemplary rather than limiting in nature. Variations and modifications to the disclosed examples may become apparent to those skilled in the art that do not necessarily depart from the essence of this disclosure. The scope of legal protection given to this disclosure can only be determined by studying the following claims.

Technology Classification (CPC): 5