Patent Publication Number: US-11643994-B2

Title: Rocket propulsion systems and associated methods

Description:
CROSS-REFERENCE TO RELATED APPLICATION 
     The present application claims priority to U.S. Provisional Application No. 62/693,829, filed on Jul. 3, 2018, and incorporated herein by reference. 
    
    
     TECHNICAL FIELD 
     The present technology is directed to rocket propulsion systems and associated methods. 
     BACKGROUND 
     Rockets propelled by rocket engines have been used for many years to launch payloads into Earth orbit and beyond. Persistent challenges associated with such systems include reducing system weight and complexity, providing adequate cooling to the rocket engine components, providing for efficient combustion, and providing efficient thrust over altitudes ranging from sea level to the vacuum of space. Aspects of the presently disclosed technology are directed toward addressing, singly and/or in combination, the foregoing challenges. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG.  1    is a partially schematic, top rear isometric view of a representative rocket-propelled aerospace vehicle configured in accordance with embodiments of the present technology. 
         FIG.  2    is a partially schematic, side view of elements of a rocket propulsion system configured in accordance with embodiments of the present technology. 
         FIG.  3 A  is a partially schematic, cross-sectional view of a conventional rocket combustion chamber configured in accordance with the prior art. 
         FIG.  3 B  is a partially schematic, cross-sectional view of a combustion chamber cooled in accordance with embodiments of the present technology. 
         FIGS.  4 A and  4 B  are partially schematic, cross-sectional views of a propellant injector configured in accordance with embodiments of the present technology. 
         FIGS.  5 A and  5 B  are partially schematic, cross-sectional illustrations of a representative injector shown in a closed position ( FIG.  5 A ) and an open position ( FIG.  5 B ), in accordance with embodiments of the present technology. 
         FIG.  6 A  is a partially schematic cross-sectional view of a representative injector shown in an open position in accordance with embodiments of the present technology. 
         FIG.  6 B  is an enlarged cross-sectional view of a portion of the injector and combustion chamber shown in  FIG.  6 A . 
         FIG.  7 A  is a partially schematic illustration of a representative propulsion system configured in accordance with embodiments of the present technology. 
         FIGS.  7 B and  7 C  illustrate representative injection ports configured in accordance with embodiments of the present technology. 
         FIGS.  7 D and  7 E  schematically illustrate a representative nozzle with fluid injection ports inactive ( FIG.  7 D ) and active ( FIG.  7 E ) in accordance with embodiments of the present technology. 
         FIG.  8    is a partially schematic, cross-sectional illustration of a fuel-cooled gas generator configured in accordance with embodiments of the present technology. 
         FIGS.  9 A and  9 B  illustrate a representative oxidizer tank and associated pressure curves as a function of time during operation in accordance with embodiments of the present technology. 
         FIGS.  10 A and  10 B  illustrate a representative fuel tank and associated pressure curves as a function of time during operation in accordance with embodiments of the present technology. 
     
    
    
     DETAILED DESCRIPTION 
     Several embodiments of the present technology are directed to rocket propulsion systems and associated methods. In representative embodiments, the propulsion system can include an injector that directs a cooling flow of rocket fuel along the walls of a spherical combustion chamber, without the need for ducts at the combustion chamber to convey the cooling flow. Accordingly, the combustion chamber can be lighter and less complex than conventional combustion chambers. The injector can be configured to deliver not only the cooling flow, but the main propellant flow into the combustion chamber, and can sequence the flows to enhance cooling. Cooling flows can also be used to cool a gas generator, which provides power to corresponding fuel and oxidizer pumps for delivering these propellant constituents into the combustion chamber. The resulting exhaust flow, exiting the combustion chamber through a nozzle, can be controlled using fluid injection to account for varying external pressures as a function of altitude. The tanks that contain the propellant constituents, and associated valving systems, can be configured to constrain the pressure differentials to which the tanks are subjected, so as to reduce the loads on the tanks, and therefore the weight of the tanks, as well as allow the tanks to have shapes that conform more closely to the contours of the vehicle. 
     Specific details of several embodiments of the disclosed technology are described below with reference to particular, representative configurations. The disclosed technology can be practiced in accordance with representative rocket configurations shown herein, and/or with other rocket systems. Although the following disclosure sets forth several representative embodiments of different aspects of the disclosed technology, some embodiments of the technology can have configurations and/or components different than those described in this section. Accordingly, the present technology can include some embodiments with additional elements, and/or without several of the elements described below with reference to  FIGS.  1 - 10 B . 
     Several embodiments of the technology described below may take the form of computer- or controller-executable instructions, including routines executed by a programmable computer or controller. Those skilled in the relevant art will appreciate that the technology can be practiced on computer/controller systems other than those shown and described below. The technology can be embodied in a special-purpose computer, controller, or data processor that is specifically programmed, configured or constructed to perform one or more of the computer-executable instructions described below. Accordingly, the terms “computer” and “controller” as generally used herein refer to any data processor and can include Internet appliances and hand-held devices (including palm-top computers, wearable computers, cellular or mobile phones, multi-processor systems, processor-based or programmable consumer electronics, network computers, mini computers and the like). Information handled by these computers can be presented at any suitable display medium, including a liquid crystal display. 
     The technology can also be practiced in distributed environments, where tasks or modules are performed by remote processing devices that are linked through a communications network. In a distributed computing environment, program modules or subroutines may be located within local and remote memory storage devices. Aspects of the technology described below may be stored or distributed on computer-readable media, including magnetic or optically readable or removable computer disks, as well as distributed electronically over networks. Data structures and transmissions of data particular to aspects of the technology are also encompassed within the scope of the embodiments of the technology. 
       FIG.  1    is a partially schematic, top rear isometric view of an aerospace vehicle  100  (which can also be referred to as a spaceplane) configured in accordance with embodiments of the present technology. In the illustrated embodiment, the vehicle  100  is an HTHL/SSTO (Horizontal Take-off, Horizontal Landing/Single Stage to Orbit) vehicle having a pair of highly swept wings  104  (identified individually as a left wing  104   a  and a right wing  104   b ) extending outwardly from a fuselage  102  to provide lift during flight in the Earth&#39;s atmosphere. The trailing edge portion of each of the wings  104   a, b  includes a corresponding elevon  106   a, b  for vehicle pitch and roll control. Additionally, the vehicle  100  includes a pair of vertical stabilizers  105  (identified individually as a left vertical stabilizer  105   a  and a right vertical stabilizer  105   b ) having corresponding rudders  108   a, b  on trailing edge portions thereof for providing the vehicle  100  with yaw control. The fuselage  102  can include a door  107   a  in an upper portion of a crew cabin  101  for crew ingress and egress. Additionally, in the illustrated embodiment the forward portion of the fuselage  102  includes a movable hatch  107   b  for providing access to a docking port (not shown in  FIG.  1   ) for docking the vehicle  100  with an on-orbit station, such as the ISS (International Space Station), and enabling human and/or cargo movement therebetween. 
     As shown in  FIG.  1   , in the illustrated embodiment the aft portion of the fuselage  102  carries a propulsion system  110  having one or more rocket engines  150  (three are shown in  FIG.  1   ; identified individually as first, second, and third rocket engines  150   a - c , respectively). Each of the engines  150   a - c  has a corresponding nozzle  160  positioned generally proximate the trailing edge portion of the wings  104   a, b  between the vertical stabilizers  105   a, b . In some embodiments the rocket engines  150   a - c  are configured to burn liquid oxygen (LOX) and jet fuel, as propellants. The jet fuel can include common kerosene-types of aviation fuel designed for use in aircraft powered by gas-turbine engines including, for example, “Jet-A.” Additionally, in some embodiments the engines  150   a - c  can include dual area ratio nozzles with injection ports for tripping the exhaust flow. When the injection ports are inactive, the exhaust flow occupies the entire cross-sectional area at the exit plane of the nozzle  160 , producing a first effective nozzle area. The area ratio of the first effective nozzle area to the area at the nozzle throat can be relatively large, which is suitable for high altitude performance. For example, the area ratio can be about 60:1 in some embodiments. When the injection ports are activated, the flow from the injection ports trips the exhaust flow and produces a shockwave that limits the effective flow over of the nozzle  160  to a smaller, second effective nozzle area. The area ratio of the second effective nozzle area to the area at the nozzle throat can be relatively small, for example, about 33:1. Accordingly, since the nozzle flow is typically over-expanded at low altitude (as a compromise to improve high altitude performance), the nozzle exit area reduction provided by the tripped exhaust flow can improve nozzle efficiency at low altitude. 
     The vehicle  100  can further include orbital maneuvering system (OMS) engines  109  (identified individually as a first OMS engine  109   a  and a second OMS engine  109   b ) having nozzles positioned just above the nozzles for the main engines  150   a - c . In some embodiments, the OMS engines  109  can be bipropellant rocket engines that use LOX and compressed natural gas (CNG; consisting mostly of methane). The use of LOX and CNG provides a gas-gas propellant solution that can be used in a blowdown system that relies on gas pressure to drive the propellants into the OMS engines  109 . The OMS engines provide steering and directional control when the vehicle  100  is in space, and can enable the vehicle  100  to reorient in space for deorbiting and reentry into the Earth&#39;s atmosphere. Although the illustrated embodiment of the vehicle  100  includes three main engines  150  that use LOX and Jet-A as propellants, the technology disclosed herein is not limited to any particular number of engines or any particular types of propellants. Accordingly, it will be understood that vehicles configured in accordance with the present technology can include more or fewer engines using other types of propellants (e.g., LOX and refined petroleum (e.g., RP-1), LOX/liquid hydrogen, LOX/CNG, etc.) consistent with the present disclosure. 
     The vehicle  100  can include a controller  190  having one or more processors  191  that can control various operations and functions of the vehicle  100  in accordance with computer-readable instructions stored on one or more system memories  192 . The controller  190  can receive inputs  113  and issue outputs  115 . By way of example, the inputs  113  can include control signals and commands from, e.g., ground systems, the crew, etc.; flight parameters such as airspeed and/or ground speed, altitude, dynamic pressure, temperature, etc.; engine operating parameters; propellant parameters; vehicle positional and directional information; etc. The outputs  115  can include commands directing vehicle operation, including control surface operation via associated valves, actuators, and/or other components; engine operation including start, stop, and throttle settings; data and telemetry transmissions; etc. The processor  191  can include any logic processing unit, such as one or more central processing units (CPUs), digital signal processors (DSPs), accelerated processing units (APUs), application-specific integrated circuits (ASICs), etc. The processor  191  may be a single processing unit or multiple processing units distributed across multiple systems and/or subsystems of the vehicle  100 . The processor  191  is operably connected to the memory  192  and may be operably connected to various systems of the vehicle  100  to transmit instructions and/or receive input therefrom. The memory  192  can include read-only memory (ROM), random-access memory (RAM), and/or other storage devices that store executable applications, test software, databases, and/or other software required to, for example, control, or at least partially control, the flight, propulsion, power, avionics, telemetry, environmental, and/or other systems of the vehicle  100  in accordance with the methods described herein, and enable the vehicle  100 , its systems and occupants to communicate and/or exchange data and information with remote computers (e.g., computers on Earth and/or in orbit) and/or other devices. 
     In some embodiments, the vehicle  100  includes all of the systems necessary for implementing the mission sequences described herein. Such systems can include, for example, a communications system  193   a  for, e.g., wireless communications (including crew communications, digital communications between processing devices, etc.) between the vehicle  100  and, e.g., ground control, ground stations, orbiting stations, etc. The communication system  193   a  can include, for example, wireless transceivers, antennae, etc. for broadcasting transmissions to, and receiving transmissions from, remote locations. The vehicle systems can also include an electrical power and distribution system  193   b ; a navigation system  193   c ; a flight controls system  193   d  for affecting actuation of the vehicle control surfaces, engine throttles, landing gear, etc.; avionics  193   e ; a hydraulic system  193   f  for, e.g., control surface and landing gear actuation; and an environmental control system  193   g  for maintaining, e.g., air conditioning, etc. for human occupancy. The foregoing systems are non-exclusive, and it will be understood that some embodiments of the vehicle  100  can include other control and operating systems, while other embodiments of the vehicle  100  may not include one or more of these systems. Further details of the representative vehicle, including launching and landing operations, are described in co-pending PCT Application No. PCT/US2019/034003, filed on May 24, 2019, and incorporated herein by reference. 
     Elements of the propulsion system  110  are further illustrated in  FIG.  2   . The propulsion system  110  can include one or more main engines  150  (one is shown in  FIG.  2   ), each of which can in turn include a combustion chamber  151  coupled to a nozzle  160 . Propellant (e.g., fuel and oxidizer) is provided by a propellant supply system  170 . The propellant supply system  170  can include a gas generator  171  which powers a turbine  172 , which, in turn, powers a fuel pump  173  and an oxidizer pump  174 . The fuel pump  173  receives fuel from a fuel supply line  114  and supplies the fuel to a propellant injector  120  via a pressurized fuel line  116 . The oxidizer pump  174  receives oxidizer from an oxidizer supply line  115  and supplies the oxidizer to the propellant injector  120  via a pressurized oxidizer line  117 . The propellant injector  120  (not clearly visible in  FIG.  2    but described in further detail later) directs the propellants into the combustion chamber  151 . The resulting exhaust products are directed through the nozzle  160  to provide thrust. Exhaust from the gas generator  171 , which is typically cooler than the exhaust from the main engine nozzle  160 , is directed to an exhaust injection manifold  161  via a turbine exhaust duct  175 , to cool the walls of the nozzle  160 . 
     In representative embodiments, the propellant includes Jet-A and the oxidizer includes liquid oxygen (LOX). In other embodiments, the engine  150  can burn other suitable propellants and/or propellant combinations. The operation of the propulsion system  110  is controlled by the automated (e.g., autonomous or semi-autonomous) controller  190  ( FIG.  1   ). 
       FIG.  3 A  is a partially schematic, cross-sectional illustration of a conventional combustion chamber  51  and an associated nozzle  60 , configured in accordance with the prior art. The combustion chamber  51  includes a propellant injector  20  that provides fuel and oxidizer to the combustion chamber  51 . The combustion chamber  51  further includes chamber walls  52 , that define a partially cylindrical central section  12  and partially hemispherical sections at each end. The chamber walls  52  are cooled by a fuel cooling flow  53  passing through associated cooling ducts  11  in the chamber walls  52 . 
     One drawback associated with the cooling arrangement described above with reference to  FIG.  3 A  is that the narrow cooling channels or ducts  11  add complexity to the manufacturing process for producing the combustion chamber  51 . Another drawback associated with the foregoing cooling arrangement is that the cooling ducts  11  add weight. Still another drawback with the foregoing cooling arrangement is that the narrow cooling channels or ducts  11  can become clogged, for example, as a result of coking in the high temperature environment of the combustion chamber  51 . Aspects of the presently disclosed technology can address, in whole or in part, the foregoing drawbacks. 
       FIG.  3 B  is a partially schematic, cross-sectional illustration of a combustion chamber  151  configured in accordance with representative embodiments of the present technology. In particular, the combustion chamber  151  can include chamber walls  152  that are spherical or nearly spherical in shape, rather than walls that include a cylindrical section  12 , as shown in  FIG.  3 A . In addition, the propellant injector  120  can direct a fuel cooling flow  153  outwardly along the curved chamber walls  152 , without the need for a cooling duct. Instead, the cooling flow  153  is delivered with sufficient momentum that it remains attached (in a fluid dynamic sense) to the chamber walls  152 , as a result of centrifugal forces. By eliminating the need for cooling ducts, this arrangement can reduce the manufacturing complexity and weight of the combustion chamber  151 . 
     In operation, the fuel cooling flow  153  can be delivered into the combustion chamber  151  at a high rate of speed (e.g., approximately 330 feet per second) and a force of approximately 3,570 g&#39;s, where one g is equal to the force of gravity at the Earth&#39;s surface. With this force, the fuel cooling flow  153  can remain attached to the chamber wall  152  throughout most of, or the entirety of, the combustion chamber  151 , as the cooling flow  153  and combustion products proceed to the throat  164  of the nozzle  160 . The thickness of the layer of fuel cooling flow  153  may change as the fuel cooling flow travels around the circumference of the combustion chamber  151 . For example, the layer thickness can decrease as the flow travels outwardly from the injector  120  toward the “equator” of the combustion chamber  151  and increase as the flow travels inwardly toward the throat  164 . In addition, as the fuel cooling flow  153  heats up, some of the flow may combust. However, in particular embodiments, the flow does not combust, and in fact passes through the throat  164  to provide further cooling to the downstream portions of the nozzle  160 . For example, the ratio of fuel-to-oxidizer in the combustion chamber  151  can be fuel-rich, so that the amount of oxidizer is insufficient to combust the cooling flow  153 , while still being sufficient to combust some or all of the main flow of fuel into the combustion chamber  151 . 
     In some embodiments, the chamber wall  152  has an entirely spherical shape e.g., up to the point at which the chamber wall  152  transitions to the throat  164 . It is expected that, with this shape, the fuel cooling flow  153  will not separate from the chamber wall  152  as it travels from the propellant injector  120  to the throat  164 . In some arrangements, the chamber wall  152  may not be entirely spherical, and may include non-spherical portions or contours. However, in such instances, the non-spherical portions are sized so as to be short enough not to cause the fuel cooling flow  153  to separate as it travels along the chamber wall  152 . 
     In addition to directing the fuel cooling flow  153 , the propellant injector  120  directs a main fuel flow  154  and a main oxidizer flow  155  into the combustion chamber  151  where the flows mix, combust (e.g., in a combustion zone within the combustion chamber), and exit through the nozzle  160 . Further details of a representative propellant injector  120  are described below with reference to  FIGS.  4 A- 6 B . 
       FIG.  4 A  is a partially schematic, cross-sectional view of a representative propellant injector  120  (e.g., a coaxial propellant injector), configured in accordance with embodiments of the present technology. The propellant injector  120  includes an injector housing  121  that is attached to, or integrally formed with, the combustion chamber  151 . The propellant injector  120  further includes an injector body  122  which carries a fuel distributor, e.g., a fuel metering ring or spray ring  123  and an oxidizer distributor  124 , e.g., a pintle tip  124   a . A fuel piston  128  and oxidizer piston  129  direct fuel and oxidizer, respectively, to the fuel metering ring  123  and to the oxidizer distributor  124 , respectively. The oxidizer piston  129  has an annular, cylindrical shape, with a channel extending upwardly within the cylindrical wall from a downwardly-facing open end. The fuel piston  128  has an annular, cylindrical shape and fits within the channel of the oxidizer piston  129 , as will be described in further detail below with reference to  FIGS.  5 A and  5 B . The two pistons  128 ,  129  can move relative to each other to coordinate the order in which fuel and oxidizer are delivered into the combustion chamber  151 . The fuel is provided to the propellant injector  120  via a fuel manifold  126  and fuel inlet ports  127 . The oxidizer is provided to the propellant injector  120  via an oxidizer inlet  125 . 
       FIG.  4 B  is an enlarged, partially schematic, cut-away illustration of the injector body  122  and internal components. As shown in  FIG.  4 B , the pintle tip  124   a  can be attached to the injector body  122  at a threaded connection (or other suitable connection)  133 . The pintle tip  124   a  includes a main oxidizer outlet (e.g., including oxidizer outlet orifices)  132  that distributes and directs the flow of oxidizer into the combustion chamber  151  ( FIG.  4 A ). The fuel metering ring  123  includes fuel cooling flow orifices  131  that direct the fuel cooling flow along the walls of the combustion chamber  151 , as described above with reference to  FIG.  3 B . The fuel metering ring  123  further includes a main fuel outlet  130  that directs the main flow of fuel into the combustion chamber  151 . The fuel piston  128  opens the flow of fuel to the fuel metering ring  123 , and the oxidizer piston  129  opens the flow of oxidizer from the pintle tip  124   a . A piston actuator port  136  provides an actuating fluid  137  (e.g., nitrogen or another inert fluid) that keeps the pistons  128 ,  129  in a normally closed position. When the fluid pressure at the piston actuator port  136  is reduced, the pistons  128 ,  129  open to the positions shown in  FIGS.  4 A and  4 B , as will be described in further detail below with reference to  FIGS.  5 A and  5 B . 
       FIG.  5 A  is a partially schematic, cross-sectional illustration of the injector  120  in a closed position or configuration. In this position, the pressure provided by the actuating fluid  137  forces the fuel piston  128  downwardly so that it engages with the fuel metering ring  123  and forms a metal-to-metal fuel seal  142 . For example, the fuel piston  128  can be formed from stainless steel and the fuel metering ring  123  can be formed from a softer metal (e.g., copper and/or copper alloys, including glidcop) which deforms slightly to form the fuel seal  142 . Accordingly, while fuel  140  may pass through the circumferentially-spaced fuel inlet ports  127  (also shown in  FIG.  4 B ), the fuel piston  128  and the fuel seal  142  prevent the fuel  140  from flowing to and through the fuel cooling flow orifices  131  and the main fuel outlet  130 . 
     The fuel piston  128  includes fuel piston seals  138  that slidably engage with an internal cavity of the oxidizer piston  129 . When the fuel piston  128  is forced downwardly to its closed position (e.g., under the force of the actuating fluid  137 ), as shown in  FIG.  5 A , it bears against an upwardly facing internal surface  148   a  of the oxidizer piston  129  and forces the oxidizer piston  129  downwardly against the pintle tip  124   a  so that the oxidizer piston  129  forms a metal-to-metal oxidizer seal  143  with the pintle tip  124   a  of the oxidizer distributor  124 . The oxidizer piston  129  can be formed from stainless steel and/or another suitable material, and the pintle tip  124   a  can be formed from glidcop or another suitable (and softer) high temperature material. The oxidizer piston  129  includes oxidizer piston seals  139  that slidably seal the oxidizer piston  129  against the internal surfaces of the injector body  122 . The injector body  122  can include one or more insulating channels  134  to prevent cryogenic oxidizer from freezing the fuel as it flows from the fuel inlet  140  into the fuel manifold  126  through the fuel inlet ports  127  and out of the fuel metering ring main fuel outlet  130  and fuel cooling flow orifices  131 . 
     When the pressure provided by the actuating fluid  137  is reduced, the force of the fuel  140  on the outer ring of the oxidizer piston  129  forces the oxidizer piston upward as indicated by arrow B, which lifts the fuel piston  128  as it bears against the upwardly facing internal surface  148   a , as indicated by arrow A. When the oxidizer piston  129  lifts, the oxidizer seal  143  unseats, and the oxidizer  141  flows outwardly through the oxidizer outlet orifices  132  in the pintle tip  124   a  so that oxidizer enters the chamber ahead of the fuel. As the fuel piston  128  moves upwardly, the fuel  140  flows to the fuel cooling flow orifices  131  and the main fuel outlet  130 . When the oxidizer piston  129  reaches its maximum upward position, the fuel piston  128  continues to move upwardly until it engages a downwardly facing surface  148   b  of the oxidizer piston  129 , at which point the fuel inlet ports  127  (visible in  FIGS.  4 A,  4 B and  5 B ) are exposed to allow fuel to flow therethrough. 
     Because the oxidizer piston  129  opens before the fuel piston  128  opens, the combustion process in the combustion chamber  151  can be initiated in a smoother manner. Accordingly, this approach can improve the combustion chamber performance. 
       FIG.  5 B  illustrates the propellant injector  120  in the fully open position. In this position, both the fuel piston  128  and the oxidizer piston  129  have lifted, allowing the fuel cooling flow  153 , the main fuel flow  154 , and the main oxidizer flow  155  to enter the combustion chamber  151 . As described above, the fuel cooling flow  153  proceeds along the chamber wall  152  under centrifugal force to cool the chamber wall  152 . The main oxidizer flow  155  is directed radially outwardly (e.g., generally in the form of a spray), and the main fuel flow  154  is directed downwardly (e.g., generally in the form of an annular cylinder) through the main oxidizer flow  155 . The main fuel flow  154  and the main oxidizer flow  155  mix and travel downwardly and outwardly into the combustion chamber  151 , where they are ignited by igniters  158  (shown schematically in  FIG.  5 B ), and combusted. 
       FIG.  6 A  is an additional illustration of the fuel and oxidizer flows when the propellant injector  120  is in the open position. As is also shown in  FIG.  6 A , a fuel pressure port  144  is used to measure the fuel pressure in the fuel manifold  126 , and a combustion chamber pressure port  156  is used to measure the static pressure in the combustion chamber  151 . 
       FIG.  6 B  is an enlarged illustration of the region around the combustion chamber pressure port  156 . The combustion chamber pressure port  156  is connected to a combustion chamber port channel  157 , which communicates with the interior of the combustion chamber  151 , e.g., via a small gap  145 . The fuel cooling flow  153  is directed tangentially along the chamber wall  152 , transverse to the gap  145 . The fuel cooling flow  153  exits through the fuel cooling flow orifices  131 , which receive fuel from a cooling flow capture slot  146 . Accordingly, a small portion of the main fuel flow  154  is diverted through the fuel cooling flow orifices  131 . 
       FIG.  7 A  is a schematic illustration of the representative propulsion system  110 , illustrating the nozzle  160 , which has an exit  167  and an injection port manifold  162 .  FIGS.  7 B and  7 C  illustrate further details of the injection port manifold  162  (which is positioned around the circumference of the nozzle  160 ), and associated injection ports. In particular,  FIG.  7 B  illustrates representative injection ports  163  that are positioned around the circumference of the nozzle  160  and receive fluid from the injection port manifold  162 .  FIG.  7 C  indicates that the injection port manifold  162  and associated injection ports  163  can be located upstream and/or downstream of the axial position shown in  FIG.  7 A , in addition to, or in lieu of the position shown in  FIG.  7 A . 
     In any of the embodiments described above with reference to  FIGS.  7 A- 7 C , the injection ports  163  can operate (e.g., under the control of the controller  190 , described above with reference to  FIG.  1   ) in the manner shown and described below with reference to  FIG.  7 D  and  FIG.  7 E . Referring first to  FIG.  7 D , when the injection ports  163  are inactive, the exhaust flow occupies the entire cross-sectional exit area of the nozzle  160 , producing a first effective nozzle area  165   a . The area ratio of the first effective nozzle area  165   a  to the area at the nozzle throat  164  can be relatively large, which is suitable for high altitude performance. For example, the area ratio can be about 60:1. 
     Referring next to  FIG.  7 E , when the injection ports  163  are activated, the flow from the injection ports  163  is directed transverse to the inner wall of the nozzle and transverse to the flow direction of exhaust products passing through the nozzle. Accordingly, the injected flow trips the exhaust flow and produces a shockwave  166  that limits the effective flow over of the nozzle  160  to a smaller, second effective nozzle area  165   b . The area ratio of the second effective nozzle area  165   b  to the area at the nozzle throat  164  can be significantly lower than what is shown in  FIG.  7 D , for example, an area ratio of 33:1. Accordingly, since the nozzle flow is typically overexpanded at low altitude (as a compromise to improve high altitude performance), absent the flow area reduction effect shown in  FIG.  7 E , the nozzle exit area reduction provided by the tripped exhaust flow can improve the nozzle efficiency at low altitude. The result is a dual area ratio engine, which can have at least two effective area ratios as shown in  FIGS.  7 D and  7 E . By varying the flow injection location (as described above with reference to  FIG.  7 C ), the area ratio can be varied over more than two values. 
     In any of the foregoing embodiments, the fluid injected into the nozzle can be inert (e.g., obtained from an on-board nitrogen bottle, or other suitable source) or, in some embodiments, can be combustible. Combustible injected fuel can provide additional thrust, in addition to providing the effective area reduction effect described above. 
       FIG.  8    is a partially schematic, cross-sectional illustration of a representative gas generator  171  configured in accordance with embodiments of the present technology. The gas generator  171  can include a combustion chamber  181  that receives fuel and oxidizer via an injector  180 , which in turn receives fuel via a main fuel inlet  176  and oxidizer via an oxidizer inlet  179 . The combustion chamber  181  can be cooled via a coolant channel  178  that receives fuel via a fuel coolant inlet  177 . The resulting coolant flow  183  is directed around the combustion chamber  181  to cool the combustion chamber  181 . The exiting coolant flow  183  can mix with the exiting exhaust products  182  to produce an exhaust products/coolant flow mixture  184  that exits the gas generator  171 . This flow can then be introduced into the main engine nozzle  160  via the turbine exhaust duct  175  and exhaust injection manifold  161 , shown in  FIGS.  2  and  7 A . Accordingly, the flow mixture  184  can be fuel-rich, with a significant amount of unburned fuel, to cool the main engine nozzle  160 . 
     The propulsion system  110  shown in  FIG.  2    receives fuel and oxidizer from corresponding fuel and oxidizer tanks.  FIGS.  9 A- 10 B  describe representative fuel and oxidizer tanks, and techniques for managing the pressures within the tanks. 
       FIG.  9 A  illustrates a representative oxidizer (or oxidant) tank  990  that supplies oxidizer (e.g., liquid oxygen) to the engine  150 . The engine  150  can be coupled to a heat exchanger  991  to heat a separate flow of unburned oxidizer, creating a flow of vaporized oxidizer  992  which is delivered into the oxidizer tank  990 , thus pressurizing it. An exit vent  993  is configured to vent the oxidizer tank  990  when the pressure differential across the walls of the oxidizer tank  990  exceeds a threshold value, for example, about 3.5 psi. An entry vent  994  is configured to allow air to enter the oxidizer tank  990  as the vehicle descends through the atmosphere, to prevent the oxidizer tank  990  from collapsing under atmospheric pressure. The entry vent  994  can include a desiccant  995  that prevents or restricts moisture from entering the oxidizer tank  990  as air enters the tank during reentry. 
       FIG.  9 B  is a graph illustrating several pressure values as a function of time during vehicle ascent. For example, line  901  illustrates the pressure outside the vehicle as it ascends from the Earth&#39;s surface beyond the atmosphere. Line  902  illustrates the absolute pressure within the oxidizer tank  990  as the vehicle ascends. Line  903  illustrates the gauge pressure or relative pressure within the tank as the vehicle ascends. As shown by line  903 , the relative pressure within the oxidizer tank  990  remains at about 3.5 psig throughout the ascent of the vehicle. During descent, the entry vent  994  and exit vent  993  can operate to maintain the pressure within the oxidizer tank  990  at 3.5 psig during descent as well. The exit vent  993  and/or the entry vent  994  can respond passively to pressure changes (e.g., via springs) or can respond actively (e.g., via solenoids or other actuators coupled to pressure sensors). 
       FIG.  10 A  is a schematic illustration of a representative fuel tank  996 , which is coupled to an entry vent  994  and a check valve  997 .  FIG.  10 B  illustrates various pressure values as a function of time during the vehicle&#39;s ascent. Referring to  FIGS.  10 A and  10 B  together, line  1001  illustrates the atmospheric pressure as a function of time during vehicle ascent, line  1002  illustrates the absolute pressure within the fuel tank  996 , and line  1003  illustrates the differential pressure within the fuel tank  996 . The fuel within the fuel tank  996  is unpressurized at launch. During ascent, the check valve  997  can close at a preselected value (e.g., 9 psia, corresponding to a non-zero altitude, e.g., 13,000 feet). The pressure within the fuel tank  996  will continue to decrease as fuel is consumed, but the relative pressure within the tank does not exceed a preselected value (e.g., about 3 psig). During vehicle descent, the entry vent  994  can be configured to release at a pressure difference of about 1 psig so that the fuel tank internal pressure tracks the atmospheric pressure during descent. As described above in the context of the oxidizer tank  990 , the fuel tank valves and vents can respond actively or passively to pressure changes. 
     For purposes of illustration, the oxidizer tank  990  and the fuel tank  996  are shown as cylindrical. In particular designs, however, these propellant tanks have conformal shapes that allow them to be placed within the wings  104   a ,  104   b , the fuselage  102 , and/or other non-cylindrical structures of the vehicle  100  shown in  FIG.  1   . As described above, the propellants for the vehicle can include liquid oxygen (LOX) and Jet-A, both of which are abundant and relatively inexpensive. Jet-A has a naturally low vapor pressure across the operating conditions of the vehicle  100 , which can allow it to be stored in tanks of virtually any shape. However, oxygen is gaseous across the vehicle  100  operating conditions, and is normally a liquid only at temperatures below its natural boiling point of −297° F. Because oxygen boils at such a low temperature, it has a high vapor pressure across the vehicle operating range and, conventionally, is kept in heavy, thick-walled, rounded or cylindrical tanks capable of withstanding higher pressures (e.g., greater than 20 psig). 
     By contrast, the LOX used for vehicles in accordance with the present technology is sufficiently cold that the vapor pressure is relatively low (e.g., less than 4 psig). For example, the LOX can be cooled from its normal boiling point (NBP) of −297° F. to −320° F. (liquid nitrogen temperature) to achieve a 5% increase in density and a 9% increase in heat sink capacity. This low vapor pressure can allow the LOX tank to have a conformal shape and therefore fit within the vehicle wing. For example, the internal structure of the wing can be wetted so that the wing itself forms a LOX tank. This approach improves aerodynamic and structural efficiency, enabling the use of lighter-weight composite materials, for example, because the tanks can support flight loads without requiring a pressure-stabilized design. The ability to shape the LOX tank over a much wider variety of conformal shapes, with reduced wall thickness, allows the vehicle to carry a greater amount of LOX per unit volume, thereby improving the mass fraction of the vehicle. 
     In operation, the LOX can be stored at a low vapor pressure, e.g., in some embodiments, to produce a pressure differential across the LOX tank walls and bulkheads of less than 3 psig. The tank pressures can be maintained at 2-3.5 psig during ascent to prevent boiling, and can be vented to prevent tank buckling during descent, as described above. Before take-off, the propellant tanks are pressurized on the ground using ground-based resources. As discussed above, once aloft, the tanks can be pressurized using vaporized propellant gasses supplied from heat exchangers coupled to the main engine. Once in orbit, residual ullage gases can vent through the exit vent  993  ( FIG.  9 A ). Because the LOX is stored at such a low pressure, the overall system can include a boost pump that increases the LOX pressure (e.g., to approximately 40 psia) to meet the inlet requirements of the oxidizer pump  174  described above with reference to  FIG.  2   . The fuel, though not at as low of a pressure as the LOX, can also include a boost pump to provide consistent inlet conditions with the turbine-driven fuel pump  173  ( FIG.  2   ). Both the fuel and the oxidizer boost pumps can include an impeller/inducer combination and can have a partial emission or full emission design, with an inducer. The pumps can be driven through a shaft connected to an electric motor or a turbine. 
     As described above, the propellant system can be sized and configured to manage and reduce pressure differentials across the walls of the associated propellant tanks. This in turn can allow the tanks to be made conformal with the vehicle, which improves the structural efficiency of the vehicle, and the propellant capacity of the vehicle. 
     Other features of the vehicle described above, which can be used singly or in combination, can produce further benefits. For example, the combustion chamber cooling design, described above, can eliminate the need for the typical multitude of small cooling channels, which are susceptible to clogging, increase the weight of the vehicle, and increase the complexity of the vehicle manufacturing process. By reducing or eliminating the likelihood for the fuel to clog, embodiments of the present technology can use less refined fuel (e.g., Jet-A rather than RP-1) to improve system flexibility and reduce cost. Embodiments of the present technology that include injectors having the configurations described above can facilitate fuel cooling flow injection, in addition to providing a simple design for both the cooling flow and the main propellant flows. 
     From the foregoing, it will be appreciated that specific embodiments of the disclosed technology have been described herein for purposes of illustration, but that various modifications may be made without deviating from the technology. For example, the overall vehicle can have configurations other than those specifically illustrated in the foregoing Figures. The propellants used for the vehicle can, in some embodiments, be different than liquid oxygen and/or Jet-A. For example, the propellant can include RP-1 (assuming costs and/other factors associated with the additional refining of RP-1 are acceptable), and/or other kerosenes. In other examples, the propellant can include liquid natural gas (LNG), compressed natural gas (CNG) and/or other methane-based fuels. Certain aspects of the technology described in the context of particular embodiments may be combined or eliminated in other embodiments. For example, an individual vehicle can include any one or combination of the foregoing subsystems, depending upon the particular application. Further, while advantages associated with certain embodiments of the disclosed technology have been described in the context of those embodiments, other embodiments may also exhibit such advantages, and not all embodiments need necessarily exhibit such advantages to fall within the scope of the technology. Accordingly, the disclosure and associated technology can encompass other embodiments not expressly shown or described herein. 
     As used herein, the term “and/or”, as in “A and/or B” refers to a alone, b alone, and both a and b. To the extent any materials incorporated herein by reference conflict with the present disclosure, the present disclosure controls. The following examples provide additional embodiments of the present technology. 
     EXAMPLES 
     1. A rocket propulsion system, comprising:
         a combustion chamber having an inwardly-facing chamber wall enclosing a combustion zone, the chamber having generally spherical shape with the chamber wall exposed to the combustion zone; and   a propellant injector coupled to the combustion chamber and having at least one fuel injector nozzle positioned to direct a flow of cooling fuel outwardly to and along the inwardly-facing chamber wall.       

     2. The system of example 1, wherein the at least one fuel injector nozzle includes a spray ring, and wherein the system further comprises an engine nozzle positioned downstream of the combustion chamber to receive combustion products produced in the combustion zone. 
     3. The system of any preceding example wherein the chamber wall does not include cooling flow channels. 
     4. The system of any preceding example wherein the at least one fuel injector nozzle is positioned to direct fuel tangentially along the inwardly-facing chamber wall. 
     5. The system of any preceding example wherein the propellent injector further includes at least one fuel outlet positioned to direct a portion of fuel axially into the combustion chamber. 
     6. The system of example 5 wherein the propellant injector further includes at least one oxidizer outlet positioned to direct oxidizer outwardly into the combustion chamber, transverse to the portion of fuel directed axially into the combustion chamber. 
     7. A rocket propulsion system, comprising
         a combustion chamber; and   a propellant injector coupled to the combustion chamber, and including:
           an oxidizer inlet;   an oxidizer distributor in fluid communication with the oxidizer inlet;   a fuel inlet;   a fuel distributor in fluid communication with the fuel inlet;   a fuel piston movably positioned between the fuel inlet and the fuel distributor;   an oxidizer piston movably positioned between the oxidizer inlet and the oxidizer distributor, and positioned to open an oxidizer flow path between the oxidizer inlet and the oxidizer distributor before the fuel piston opens a fuel flow path between the fuel inlet and the fuel distributor.   
               

     8. The system of example 7 wherein the fuel piston is slidable relative to the oxidizer piston. 
     9. The system of example 7 wherein the fuel piston and the oxidizer piston each have an annular shape, and wherein the fuel piston is positioned within an annulus of the oxidizer piston. 
     10. The system of example 9, further comprising a piston actuator port in fluid communication with the fuel piston and the oxidizer piston. 
     11. The system of example 10 wherein the fuel piston is positioned to engage a first surface of the oxidizer piston to drive both the fuel piston and the oxidizer piston to respective closed positions when a pressure at the actuator port is at a first level, and wherein the fuel piston and the oxidizer piston move relative to each other to respective open positions when a pressure at the actuator port is at a second level less than the first level. 
     12. The system of example 7 wherein the oxidizer distributer includes a plurality of oxidizer outlet orifices positioned to direct a spray of oxidizer radially outwardly, and wherein the fuel distributer includes a fuel outlet positioned to direct a cylindrical flow of fuel transversely to and through the spray of oxidizer. 
     13. The system of example 7 wherein the propellant injector includes:
         cooling flow orifices positioned to direct a first portion fuel tangentially along an inner wall of the combustion chamber   a fuel flow outlet positioned to direct a second portion of fuel axially into the combustion chamber; and   an oxidizer outlet positioned to direct oxidizer outwardly into the combustion chamber, transverse to the second portion of fuel.       

     14. The system of example 7, further comprising:
         an exhaust nozzle coupled to the combustion chamber, the exhaust nozzle having a throat and an exit downstream of the throat, the exit having an exit cross-sectional area, the nozzle further including a plurality of injection ports positioned to direct an injected flow transverse to an inner wall of the nozzle; and   a controller operatively coupled to the injection ports to (a) direct injection of a fluid through the injection ports to reduce an effective cross-sectional area of the nozzle for operation at a first altitude, and (b) cease injection of the fluid at a second altitude greater than the first altitude.       

     15. A reusable space vehicle system, comprising:
         a reusable horizontal takeoff/horizontal landing (HTHL), ground-assisted single-stage-to-orbit (SSTO) spaceplane; and   a rocket propulsion system, carried by the spaceplane, and including:
           a combustion chamber having an inwardly-facing chamber wall enclosing a combustion zone, the chamber having generally spherical shape with the chamber wall exposed to the combustion zone; and   a propellant injector coupled to the combustion chamber and having at least one fuel injector nozzle positioned to direct a flow of cooling fuel outwardly along the inwardly-facing chamber wall.   
               

     16. The system of example 15 wherein the chamber wall does not include cooling flow channels. 
     17. The system of any of examples 15-16 wherein the at least one fuel injector nozzle is spaced apart from the inwardly-facing chamber wall. 
     18. The system of any of examples 15-17 wherein the propellent injector further includes:
         at least one fuel outlet positioned to direct a portion of fuel axially into the combustion chamber; and   at least one oxidizer outlet positioned to direct oxidizer outwardly into the combustion chamber, transverse to the portion of fuel directed axially into the combustion chamber.       

     19. The system of any of examples 15-18 wherein the propellant injector includes:
         an oxidizer inlet;   an oxidizer distributor in fluid communication with the oxidizer inlet;   a fuel inlet;   a fuel distributor in fluid communication with the fuel inlet, the fuel distributer including the at least one fuel injector nozzle positioned to direct a flow of cooling fuel outwardly along the inwardly-facing chamber wall;   a fuel piston movably positioned between the fuel inlet and the fuel distributor;   an oxidizer piston movably positioned between the oxidizer inlet and the oxidizer distributor, and positioned to open an oxidizer flow path between the oxidizer inlet and the oxidizer distributor before the fuel piston open a fuel flow path between the fuel inlet and the fuel distributor.       

     20. The system of example 19, further comprising
         a piston actuator port in fluid communication with the fuel piston and the oxidizer piston; and wherein   the fuel piston is positioned to engage a first surface of the oxidizer piston to drive both the fuel piston and the oxidizer piston to respective closed positions when a pressure at the actuator port is at a first level, and wherein the fuel piston and the oxidizer piston move relative to each other to respective open positions when a pressure at the actuator port is at a second level less than the first level.       

     21. The system of any of examples 15-20, further comprising:
         an oxidant tank coupled to the propellant injector;   an exit vent coupled to the oxidant tank and being configured to release oxidant from the oxidant tank when a pressure differential across a wall of the oxidant tank exceeds a threshold value; and   an entry vent coupled to the oxidant tank and positioned to permit air entry into the oxidant tank during vehicle descent.       

     22. The system of any of examples 15-21, further comprising:
         a fuel tank coupled to the propellant injector;   a check valve coupled to the fuel tank and being configured to close at a preselected pressure corresponding to a non-zero altitude; and   an entry vent coupled to the fuel tank and configured to release at a selected pressure differential to permit air entry into the fuel tank during vehicle descent.