Abstract:
A small satellite propulsion system using a gaseous oxidizer and a gaseous fuel as primary propellants with a liquid as a film coolant for the inner surface of the rocket motor. The gaseous fuel is also used as a pressurant for the coolant and as a cold gas propellant for attitude control system (hereinafter “ACS”) thrusters. The oxidizer, fuel, and coolant tanks, as well as most valves and plumbing, are integrated into a single core unit along with the rocket motor, rocket motor plumbing, and safety valves. Attitude control thrusters may be remotely located with plumbing to the fuel tank. The core unit is four inches high and less than four inches deep and wide. The small satellite propulsion system uses no pyrotechnics and no hazardous toxic materials.

Description:
RELATIONSHIP TO OTHER APPLICATIONS 
       [0001]    The present application claims the benefit of U.S. provisional patent application Ser. No. 62/103,204 filed Jan. 14, 2015 to the same inventors. 
     
    
     STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH AND DEVELOPMENT 
       [0002]    This invention was made with government support under Phase I SBIR contract NNX14CP20P awarded by NASA. The government has certain rights in the invention. 
     
    
     FIELD OF ART 
       [0003]    The present invention relates to a propulsion system for a small satellite, such as a three-unit or six-unit (3U, 6U, or larger) cubesat. The present invention more particularly relates to a propulsion system for a small satellite that does not use any pyrotechnics or any hazardous toxic materials. 
       BACKGROUND OF THE INVENTION 
       [0004]    Cubesats are a variety of small satellite having a standard size and shape. While the illustrated embodiment below is oriented toward a cubesat, other types of small satellites may also use the present invention. The inventive small satellite propulsion system strikes a balance between high performance (which typically mandates use of toxic propellants such as hydrazine, peroxide or ammonia) and safety mandates (that limit use of pressurized chemical propellants), and seeks to overcome technical and programmatic constraints for onboard cubesat propulsion. The technical challenge is to realize such propulsion in a suitably small volume, especially once tanks, thrusters, valves, controls, etc. are added. The programmatic challenge in cubesat propulsion is also complicated by design guidelines that presently prohibit pyrotechnics, pressure vessels over 1.2 atmospheres, use of hazardous materials, and storage of more than 100 W-Hrs of chemical energy. 
       SUMMARY OF THE INVENTION 
       [0005]    Briefly described, the invention includes a small satellite propulsion system using a gaseous oxidizer and a gaseous fuel as primary propellants with a liquid as a film coolant for the inner surface of the rocket motor. The gaseous fuel is also used as a pressurant for the coolant and as a cold gas propellant for attitude control system (hereinafter “ACS”) thrusters. The oxidizer, fuel, and coolant tanks, as well as most valves and plumbing, are integrated into a single core unit along with the rocket motor, rocket motor plumbing, and safety valves. Attitude control thrusters may be remotely located with plumbing to the fuel tank. The core unit is four inches high and less than four inches deep and wide. The small satellite propulsion system uses no pyrotechnics and no hazardous toxic materials. 
         [0006]    In exemplary embodiments, the invention includes a small satellite propulsion system including: a gaseous fuel tank; a gaseous oxidizer tank; a liquid coolant tank; and a rocket motor in valve-controlled fluid communication with the gaseous fuel tank, the gaseous oxidizer tank, and the liquid coolant tank, where the rocket motor is fixed within an elongated enclosure including at least one side formed, at least in part, by the gaseous fuel tank, the gaseous oxidizer tank, and/or the liquid coolant tank. That small satellite propulsion system, including a piston in the liquid coolant tank, responsive to a pressurant on a first side of the piston to exert pressure on a liquid coolant in the liquid coolant tank when the small satellite propulsion system is fueled, filled, and activated. That small satellite propulsion system, including valve-controlled fluid communication between the gaseous fuel tank and the liquid coolant tank adapted to supply gaseous fuel as the pressurant when the small satellite propulsion system is fueled, filled, and activated. That small satellite propulsion system, including valve-controlled fluid communication between the gaseous fuel tank and an attitude control system adapted to use gaseous fuel in cold-gas attitude control system thrusters when the small satellite propulsion system is fueled, filled, and activated. That small satellite propulsion system, where the gaseous fuel includes gaseous methane, the gaseous oxidizer includes gaseous oxygen, and the liquid coolant includes ethanol and water. That small satellite propulsion system, where the liquid coolant tank contains a pressure no greater than 1.2 atmospheres when the small satellite propulsion system is fueled and filled. That small satellite propulsion system, where the gaseous fuel tank, the gaseous oxidizer tank, and the liquid coolant tank together contain no more than 100 Watt-hours of chemical energy when the small satellite propulsion system is fueled and filled. That small satellite propulsion system, including no pyrotechnics. That small satellite propulsion system, including no hazardous toxic materials when the gaseous fuel tank, the gaseous oxidizer tank, and the liquid coolant tank are filled. That small satellite propulsion system, where the valve-controlled fluid communication includes: gas valve-controlled fluid communication with the rocket motor, further including: first and second isolation valves in fluid communication with the respective GOX and GCH4 tanks; first and second gas regulation valves in fluid communication with the first and second respective isolation valves; first and second run valves in fluid communication with the first and second respective gas regulation valves; first and second check valves in fluid communication with the first and second respective run valves; first and second flow venturies in fluid communication with the first and second respective check valves; and a rocket motor injector in fluid communication with the first and second flow venturies; gas valve-controlled fluid communication with the attitude control system thrusters, further including: the second isolation valve; the second gas regulation valve; four thruster run valves in fluid communication with the second gas regulation valve; and four thrusters in fluid communication with the respective four thruster run valves; gas valve-controlled fluid communication with the coolant tank, further including: the second isolation valve; the second gas regulation valve; a pressurant run valve in fluid communication with the second gas regulation valve; and a pressurant side of the piston in the coolant tank in fluid communication with the pressurant run valve; and liquid valve-controlled fluid communication, further including a third isolation valve in fluid communication with the coolant tank; a third run valve in fluid communication with the third isolation valve; a third check valve in fluid communication with the third run valve; and the rocket motor injector in fluid communication with the third check valve. 
         [0007]    A small satellite propulsion system including: a gaseous fuel tank; a gaseous oxidizer tank; a liquid coolant tank; a rocket motor in valve-controlled fluid communication with the gaseous fuel tank, the gaseous oxidizer tank, and the liquid coolant tank, where the rocket motor is fixed within an elongated enclosure including at least one side formed, at least in part, by the gaseous fuel tank, the gaseous oxidizer tank, and/or the liquid coolant tank; a piston in the liquid coolant tank, responsive to a pressurant on a first side of the piston to exert pressure on a liquid coolant in the liquid coolant tank when the small satellite propulsion system is fueled, filled, and activated; and valve-controlled fluid communication between the gaseous fuel tank and the liquid coolant tank adapted to supply gaseous fuel as the pressurant when the small satellite propulsion system is fueled, filled, and activated. That small satellite propulsion system, including valve-controlled fluid communication between the gaseous fuel tank and an attitude control system adapted to use gaseous fuel in cold-gas attitude control system thrusters when the small satellite propulsion system is fueled, filled, and activated. That small satellite propulsion system, where the gaseous fuel tank, the gaseous oxidizer tank, and the liquid coolant tank each contains a pressure no greater than 1.2 atmospheres when the small satellite propulsion system is fueled and filled. That small satellite propulsion system, where the gaseous fuel includes gaseous methane, the gaseous oxidizer includes gaseous oxygen, and the liquid coolant includes ethanol and water. That small satellite propulsion system, where the gaseous fuel tank, the gaseous oxidizer tank, and the liquid coolant tank together contain no more than 100 Watt-hours of chemical energy when fueled and filled. That small satellite propulsion system, including: no pyrotechnics, no hazardous toxic materials when the gaseous fuel tank, the gaseous oxidizer tank, and the liquid coolant tank are fueled and filled. That small satellite propulsion system, where the valve-controlled fluid communication includes at least one of: gas valve-controlled fluid communication, further including: first and second isolation valves in fluid communication with the respective GOX and GCH4 tanks; first and second gas regulation valves in fluid communication with the first and second respective isolation valves; first and second run valves in fluid communication with the first and second gas regulation valves; first and second check valves in fluid communication with the first and second run valves; first and second flow venturies in fluid communication with the first and second check valves; and a rocket motor injector in fluid communication with the first and second flow venturies; and a third isolation valve in fluid communication with the coolant tank; a third run valve in fluid communication with the third isolation valve; a third check valve in fluid communication with the third run valve; and the rocket motor injector in fluid communication with the third check valve. 
         [0008]    A small satellite propulsion system including: a gaseous fuel tank; a gaseous oxidizer tank; a liquid coolant tank; a rocket motor in valve-controlled communication with the gaseous fuel tank, the gaseous oxidizer tank, and the liquid coolant tank, where the rocket motor is fixed within an elongated enclosure including at least one side formed, at least in part, by at least one of the gaseous fuel tank, the gaseous oxidizer tank, and the liquid coolant tank; a piston in the liquid coolant tank, responsive to a pressurant on a first side of the piston to exert pressure on a liquid coolant in the liquid coolant tank when the small satellite propulsion system is fueled, filled, and activated; valve-controlled fluid communication between the gaseous fuel tank and the liquid coolant tank adapted to supply gaseous fuel as the pressurant when the small satellite propulsion system is fueled, filled, and activated; and valve-controlled fluid communication between the gaseous fuel tank and an attitude control system adapted to use gaseous fuel in cold-gas attitude control system thrusters when the small satellite propulsion system is fueled, filled, and activated. That small satellite propulsion system, where the valve-controlled fluid communication includes at least one of: gas valve-controlled fluid communication with the rocket motor, further including: first and second isolation valves in fluid communication with the respective GOX and GCH4 tanks; first and second gas regulation valves in fluid communication with the first and second respective isolation valves; first and second run valves in fluid communication with the first and second respective gas regulation valves; first and second check valves in fluid communication with the first and second respective run valves; first and second flow venturies in fluid communication with the first and second respective check valves; and a rocket motor injector in fluid communication with the first and second flow venturies; gas valve-controlled fluid communication with the attitude control system thrusters, further including: the second isolation valve; the second gas regulation valve; four thruster run valves in fluid communication with the second gas regulation valve; and four thrusters in fluid communication with the respective four thruster run valves; gas valve-controlled fluid communication with the coolant tank, further including: the second isolation valve; the second gas regulation valve; a pressurant run valve in fluid communication with the second gas regulation valve; and a pressurant side of the piston in the coolant tank in fluid communication with the pressurant run valve; and liquid valve-controlled fluid communication, further including a third isolation valve in fluid communication with the coolant tank; a third run valve in fluid communication with the third isolation valve; a third check valve in fluid communication with the third run valve; and the rocket motor injector in fluid communication with the third check valve. That small satellite propulsion system, where the gaseous fuel tank, the gaseous oxidizer tank, and the liquid coolant tank each contains a pressure no greater than 1.2 atmospheres when the small satellite propulsion system is fueled and filled; the gaseous fuel tank, the gaseous oxidizer tank, and the liquid coolant tank together contain no more than 100 Watt-hours of chemical energy when small satellite propulsion system is fueled and filled; small satellite propulsion system includes: no pyrotechnics; and no hazardous toxic materials. 
     
    
     
       DESCRIPTION OF THE FIGURES OF THE DRAWINGS 
         [0009]    The present invention will hereinafter be described in conjunction with the following drawing figures, wherein like numerals denote like elements, and 
           [0010]      FIG. 1  is a perspective view, with partial transparency, illustrating an exemplary embodiment of the small satellite propulsion system core unit, according to a preferred embodiment of the present invention; 
           [0011]      FIG. 2  is a front elevation cross section view illustrating an exemplary embodiment of the small satellite propulsion system core unit of  FIG. 1 , according to a preferred embodiment of the present invention; 
           [0012]      FIG. 3  is a plumbing and instrumentation diagrammatic view illustrating an exemplary embodiment of the small satellite propulsion system, according to a preferred embodiment of the present invention; 
           [0013]      FIG. 4  is a perspective view illustrating an exemplary embodiment of the rocket motor and thrust chamber test article in the small satellite propulsion system, according to a preferred embodiment of the present invention; 
           [0014]      FIG. 5  is a perspective view illustrating an exemplary embodiment of the rocket motor and thrust chamber test article in the small satellite propulsion system, according to a preferred embodiment of the present invention; 
           [0015]      FIG. 6  is an additional perspective view illustrating an exemplary embodiment of the rocket motor and the thrust chamber test article in the small satellite propulsion system, according to a preferred embodiment of the present invention; 
           [0016]      FIG. 7  is an additional perspective view illustrating an exemplary embodiment of the rocket motor and the thrust chamber test article in the small satellite propulsion system, according to a preferred embodiment of the present invention; 
           [0017]      FIG. 8  is a perspective view illustrating an exemplary structure of a second embodiment of a small satellite propulsion system, according to a preferred embodiment of the present invention; 
           [0018]      FIG. 9  is a perspective cross-sectional view illustrating the exemplary structure of the second embodiment of a small satellite propulsion system of  FIG. 8 , according to a preferred embodiment of the present invention; 
           [0019]      FIG. 10  is a top plan view illustrating the exemplary second embodiment of a small satellite propulsion system of  FIG. 8  and defining cross-sections AA and BB, according to a preferred embodiment of the present invention; 
           [0020]      FIG. 11  is an elevation cross-sectional view, through cross section AA defined in  FIG. 10 , illustrating the exemplary second embodiment of a small satellite propulsion system of  FIG. 8 , according to a preferred embodiment of the present invention; and 
           [0021]      FIG. 12  is an elevation cross-sectional view, through cross section BB defined in  FIG. 10 , illustrating the exemplary second embodiment of a small satellite propulsion system of  FIG. 8 , according to a preferred embodiment of the present invention. 
       
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
       [0022]      FIG. 1  is a perspective view, with partial transparency, illustrating an exemplary embodiment of the small satellite propulsion system  300  (see  FIG. 3 ) core unit  100 , according to a preferred embodiment of the present invention. Core unit  100  includes toroidal gaseous Oxygen (hereinafter “GOX”) tank  102  mounted on toroidal gaseous methane (hereinafter “GCH4”) tank  104  that is mounted on annular coolant tank  108 . The open centers of the toroidal tanks  102  and  104 , as well as the open center of annular coolant tank  108  houses valves and plumbing, discussed in further detail below, and the rocket motor  110 . 
         [0023]    In another embodiment of the system, to be discussed further below, the toroidal tanks may be replaced by longitudinal tanks. In other embodiments, nitrous oxide may be used as an oxidizer. In other embodiment, gaseous ethane may be used as a fuel. 
         [0024]    The coolant is a mixture of ethanol and water. The coolant is stored as an unpressurized liquid for film cooling of the chamber of the rocket motor  110 , and operates with a depressed freezing point down to −30° C. After deployment, annular coolant tank  108  is pressurized by gas from the GCH4 tank  104  via a high-density polyethylene (hereinafter “HDPE”) piston  106  to provide injection pressure to create a cooling film stream on the inner walls of rocket motor  110 . The coolant is injected though a peripheral arrangement of coolant ports in the injector of the rocket motor  110 . The coolant may also be injected into the side walls of the chamber upstream of the nozzle for transpiration cooling, or circulate through cooling channels in the walls of the thruster in a cooled-nozzle configuration for a rocket engine. COTS burst disk pressure relief valve  116  for the annular coolant tank  108  provides an ASME compliant factor of safety to allow for realization of a safe flight-ready system. Thruster cavity  118  is an elongated enclosure having at least one side  120  formed, at least in part, by at least one of the coolant tank  108 , the GCH4 tank  104 , and the GOX tank  102 . 
         [0025]    The toroidal GOX and GCH4 tanks  102  and  104  are each machined from solid 7075-T6 aluminum as two separate pieces and fastened together using O-ring seals. In other embodiments, other materials of similar functional characteristics may be used for forming the tanks  102  and  104 . Alternatively, the longitudinal tanks may be milled into a structural aluminum or other metal block. The toroidal GOX and GCH4 tanks  102  and  104  are installed as inert, sealed components with burst disk relief valves designed for a factor of safety of four (equivalent to ASME code). GOX tank  102  is filled through GOX fill valve  112 , which is preferably a quick-disconnect GOX fill valve  112 . GCH4 tank  104  is filled through GCH4 fill valve  114 , which is preferably a quick-disconnect GCH4 fill valve  114 . The core unit  100  of the illustrated embodiment is preferably four inches in height. 
         [0026]      FIG. 2  is a front elevation cross section view illustrating an exemplary embodiment of the small satellite propulsion system  300  (see  FIG. 3 ) core unit  100  of  FIG. 1 , according to a preferred embodiment of the present invention. Within the center space of the toroidal tanks  102  and  104 , an isolation valve  202  is installed above the rocket motor  110 . Isolation valves  202  (one of three is visible in this view) utilize a motor-driven screw to pierce a positive fluid isolation barrier on each of the methane and oxygen tanks  104  and  102  once the small satellite propulsion system  300  (see  FIG. 3 ) is ready to be activated in orbit. Injector  204  is mounted atop the rocket motor  110  and includes a manifold for directing coolant, GOX and GCH4 into the combustion chamber  210  of the rocket motor  110 . Also shown is a rocket motor pressure transducer  206  for the combustion chamber  210  that is used in a control system (not shown) for the small satellite propulsion system  300 . Also shown is the GOX run valve  208 . GOX run valve  208  is the operational oxidizer control valve. 
         [0027]      FIG. 3  is a plumbing and instrumentation diagrammatic view illustrating an exemplary embodiment of the small satellite propulsion system  300 , according to a preferred embodiment of the present invention. GOX tank  102  has a GOX tank temperature transducer  312  and a GOX tank pressure transducer  314  providing outputs to a control system (not shown). Main GOX line  302  is filled before launch through quick-disconnect GOX fill valve  112 , as described above. GOX relief valve  304  is preferably a burst disc GOX relief valve  304 , similar to burst disk pressures relief valve  116 , as described above. GOX isolation valve  202  is a single-use valve that opens the main GOX line  302  to the rocket motor  110  when the small satellite propulsion system  300  is deployed in outer space. GOX gas regulation valve  308  provides a regulated outlet pressure. GOX run valve  208 , discussed above, is a solenoid controlled GOX run valve  208  that operates based on the control system (not shown) sending control signals to the solenoid. GOX check valve  306  prevents back pressure from the rocket motor  110  from reaching main GOX line  302 . The output of GOX check valve  306  is a controlled flow of GOX in GOX controlled line  360  to the GOX flow venturi  372  and then to the injector  204  and thence into the combustion chamber  210  of rocket motor  110 . GOX line pressure transducer  310 , similar to GOX tank pressure transducer  314 , discussed above, provides pressure data to the control system (not shown) regarding the pressure of GOX at the injector  204 . Rocket motor pressure transducer  206  provides pressure data regarding the pressure inside of combustion chamber  210  of rocket motor  110  via rocket motor pressure transducer line  370 . 
         [0028]    Coolant is loaded into annular coolant tank  108  through quick disconnect fill valve  348  and coolant main line  350 . Annular coolant tank  108  has a burst disk pressure relief valve  116 , as discussed above, a coolant pressure transducer  346 , a coolant temperature transducer  348 , and a pressurant input line  364 . During operation, HDPE piston  106  applies pressure, during operation, to the coolant in annular coolant tank  108 . Coolant exits annular coolant tank  108  through main coolant line  350  to coolant isolation valve  352 , which is similar to GOX isolation valve  202 , discussed above. The flow through coolant isolation valve  352  goes to coolant run valve  354 , similar to GOX run valve  208  discussed above, which is controlled by control inputs to the solenoid. Coolant check valve  356 , similar to GOX check valve  306  discussed above, passes coolant unidirectionally along coolant controlled flow line  362  to the rocket motor  110  and prevents pressure from the rocket motor  110  from backing up into the annular coolant tank  108 . 
         [0029]    GCH4 tank  104  has a GCH4 temperature transducer  338  and a GCH4 pressure transducer  340  to supply inputs to the control system (not shown). GCH4 exits GCH4 tank  104  through GCH4 main line  316 . GCH4 tank  104  is filled via quick disconnect GCH4 fill valve  114 , similar to the GOX quick disconnect fill valve  112  discussed above. GCH4 tank  104  has a GCH4 burst disk relief valve  318 , similar to the GOX burst disk relief valve  304  discussed above, connected via GCH4 main line  316 . GCH4 flow from GCH4 tank  104  is through GCH4 isolation valve  322  that is similar to GOX isolation valve  202  discussed above. GCH4 gas regulation valve  324 , similar to GOX gas regulation valve  308  discussed above, supplies regulated pressure to the GCH4 pressurant line  342 , to the GCH4 ACS cold gas monopropellant line  336 , and to the rocket motor fuel line  366 . 
         [0030]    GCH4 pressurant line  342  couples to GCH4 pressurant run valve  344  (similar to GOX run valve  208 , discussed above). GCH4 pressurant run valve  344  is controlled by control system (not shown) signals to the solenoid. In operation, controlled GCH4 pressurant flow is conducted along coolant controlled pressurant input line  364  to annular coolant tank  108  to apply force to HDPE piston  106 . 
         [0031]    In the GCH4 ACS thruster assembly  330 , ACS cold gas monopropellant line  336  conducts regulated GCH4 to GCH4 ACS thrusters  334  (one of four labeled) via respective GCH4 ACS thruster run valves  332  (one of four labeled), similar to the GOX run valve  208  discussed above. The ACS thrusters  334  are not part of the core unit  100 , but are remotely located on the satellite to provide small amounts of thrust for attitude control and station keeping. In a particular embodiment, more than four ACS thrusters  334  and their respective run valves  332  may be used. 
         [0032]    Rocket motor fuel line  366  supplies pressure regulated GCH4 to GCH4 run valve  326  (similar to GOX run valve  208 , discussed above), which controls the flow of GCH4 fuel to rocket motor  110  via GCH4 check valve  328  (similar to GOX check valve  306 , discussed above), GCH4 injector  204  (similar to GOX injector  204 , discussed above), GCH4 controlled line  358  and GCH4 flow venturi  368 . 
         [0033]      FIG. 4  is a perspective view illustrating an exemplary embodiment of an exemplary rocket motor  110  and thrust chamber test article  610  of the small satellite propulsion system  300 , according to a preferred embodiment of the present invention. Rocket motor  110  is approximately one and three-quarters of an inch high in the illustrated embodiment. A top rocket motor portion  406  houses interface plug  404 , which is secured in the top rocket motor portion  406  by set screws  408  (two of four visible in this view, one of two labeled). The injector  204 , not shown in this view, holds the terminations of the rocket motor pressure transducer line  370 , the GOX controlled line  360 , the GCH4 controlled line  358 , and the coolant controlled flow line  362 . Also held in the injector  204  is a miniature spark plug igniter  402 , which ignites the GOX/GCH4 mixture in the rocket motor  110  during operation. 
         [0034]      FIG. 5  is a perspective view illustrating an exemplary embodiment of the rocket motor  110  and thrust chamber test article  610  in the small satellite propulsion system  300 , according to a preferred embodiment of the present invention. The injector  204  holds the terminations of the rocket motor pressure transducer line  370 , the GOX controlled line  360 , the GCH4 controlled line  358 , and the coolant controlled flow line  362 . Also visible is the threaded bore  504  for installing the miniature spark plug igniter  402 . US penny  502  is included for size comparison. 
         [0035]      FIG. 6  is an additional perspective view illustrating an exemplary embodiment of the rocket motor  110  and the thrust chamber test article  610  in the small satellite propulsion system  300 , according to a preferred embodiment of the present invention. Thrust chamber test article  610  is shown with four pipes and nuts and a flanged support for the plumbing. US penny  502  is included for size comparison. 
         [0036]      FIG. 7  is an additional perspective view illustrating an exemplary embodiment of the rocket motor  110  and the thrust chamber test article  610  in the small satellite propulsion system  300 , according to a preferred embodiment of the present invention. Thrust chamber test article  610  is again shown with the four pipes (the rocket motor pressure transducer data line  370 , the GOX controlled line  360 , the GCH4 controlled line  358 , and the coolant controlled flow line  362  (barely visible)) and nuts and a flanged support for the plumbing. 
         [0037]      FIG. 8  is a perspective view illustrating an exemplary structure  802  of a second embodiment of a small satellite propulsion system  800 , according to a preferred embodiment of the present invention. Structure  802  encloses linear GCH4 tanks  806  and  810 , linear GOX tanks  808  and  812 , and a thruster cavity  814 . Thruster cavity  814  is an elongated enclosure having at least one wall that is formed, at least in part, by coolant tanks  904  (see  FIG. 9 ). Top plate  816  is fastened to the structure  802 . Structure  802  has a height  804  which, in the present exemplary embodiment, is four inches. 
         [0038]      FIG. 9  is a perspective cross-sectional view illustrating the exemplary structure  802  of the second embodiment of a small satellite propulsion system  800  of  FIG. 8 , according to a preferred embodiment of the present invention. Inner tank walls  906  and  908  of linear GCH4 tanks  810  and  806 , respectively, can be seen. Linear coolant tanks,  904  (one of two labeled) is shown with GCH4-driven piston  902  (one of two labeled). Linear coolant tanks,  904  (one of two labeled) are positioned between GCH4 tank  810  and thruster cavity  814  and between GCH4 tank  806  and thruster cavity  814 . Linear coolant tanks,  904  (one of two labeled) are fluidically connected (not shown). 
         [0039]      FIG. 10  is a top plan view illustrating the exemplary second embodiment of a small satellite propulsion system  800  of  FIG. 8  and defining cross-sections AA and BB, according to a preferred embodiment of the present invention. Section AA is through GCH4 tanks  806  and  810  at two opposing corners of structure  802  and Section BB is through GOX tanks  808  and  812  at the other two opposing corners of structure  802 . 
         [0040]    GCH4 fill valve  1002  (similar to GCH4 fill valve  114 ) receives GCH4 prior to launch to fill GCH4 tanks  806  and  810  through GCH4 tank closures  1020  and  1024 , respectively. Tank  806  and tank  810  are fluidically connected via GCH4 line  1022  extending through first tank closure  1020  into second tank closure  1024 , respectively. GCH4 from GCH4 tanks  806  and  810  flows via GCH4 output line  1030  and GCH4 regulator valve  1004  into manifold  1018  when GCH4 isolation valve  1006  (similar to GCH4 isolation valve  322 ) is open. From GCH4 isolation valve  1006 , GCH4 flows to GCH4 run valve  1102  (see  FIG. 11 ) and then into injector  204  of rocket motor  110  in thruster cavity  814 . Note that the GCH4 flow illustrated in  FIG. 10  differs from the embodiment of  FIG. 3 . 
         [0041]    GOX fill valve  1012  (similar to valve  112 ) receives GOX prior to launch to fill GOX tanks  808  and  812  through GOX tank closures  1026  and  1032 , respectively. GOX tank  808  and GOX tank  812  are fluidically connected via GOX line  1028  extending through first GOX tank closure  1026  into second GOX tank closure  1032 , respectively. GOX from GOX tanks  808  and  812  flows to GOX regulator valve  1014  (similar to GOX gas regulation valve  308 ) via GOX output line  1034  and into manifold  1018  via when GOX isolation valve  1016  (similar to isolation valve  202 ) is open. From GOX isolation valve  1016 , GOX flows to GOX run valve  1202  (see  FIG. 12 ) and then into injector  204  of rocket motor  110  in thruster cavity  814   
         [0042]    Computer interface  1010  provides a connection point for a control computer communication line (not shown). The control computer, which may be one of many programs running on a single non-dedicated computer or a dedicated computational resource, operates the valves and receives status and sensor data from the small satellite propulsion system  800 . The control computer implements both pre-deployment and operational actions of the small satellite propulsion system  800 . 
         [0043]      FIG. 11  is an elevation cross-sectional view, through cross section AA defined in  FIG. 10 , illustrating the exemplary second embodiment of a small satellite propulsion system  800  of  FIG. 8 , according to a preferred embodiment of the present invention. GCH4 run valve  1102  (similar to GCH4 run valve  326 ) from the manifold  1018  into the thruster cavity  814 . Coolant run valve  1106  (similar to coolant run valve  354 ) is located in the thruster cavity  814  where it receives coolant thought the coolant tank sidewall. GCH4 flows through the controlled line from the GCH4 run valve  1102  to the injector  204 . Coolant controlled flow line  1104  (similar to coolant controlled flow line  362 ) is connected (not shown, but see  FIG. 12 ) to injector  204 . Linear GCH4 tanks  806  and  810  have respective GCH4 first and second tank closures  1020  and  1024  installed. Coolant tanks  904  (one of two labeled) are fluidically connected (not shown) for liquid coolant at the tank bottoms and for GCH4 pressure at the tank tops. 
         [0044]      FIG. 12  is an elevation cross-sectional view, through cross section BB defined in  FIG. 10 , illustrating the exemplary second embodiment of a small satellite propulsion system  800  of  FIG. 8 , according to a preferred embodiment of the present invention. GOX tanks  808  and  812  are shown with tank closures  1026  and  1032  installed, respectively. GOX run valve  1202  (similar to GOX run valve  208 ) extends from the manifold  1018  into the thruster cavity  814 , as does GCH4 run valve  1102 . Rocket motor  110  is mechanically fixed within thruster cavity  814 .