Patent Publication Number: US-11649070-B2

Title: Earth to orbit transportation system

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS INCORPORATED BY REFERENCE 
     The present application is a continuation of U.S. patent application Ser. No. 16/745,187, filed Jan. 16, 2020, and titled EARTH TO ORBIT TRANSPORTATION SYSTEM, now issued as U.S. Pat. No. 11,059,608, which is a continuation of International Patent Application No. PCT/US19/34003, filed May 24, 2019, and titled EARTH TO ORBIT TRANSPORTATION SYSTEM, which claims priority to U.S. Provisional Patent Application No. 62/676,809, filed May 25, 2018, and titled EARTH TO ORBIT TRANSPORTATION SYSTEM, each of which is incorporated herein by reference in its entirety. 
    
    
     TECHNICAL FIELD 
     The present disclosure is generally related to vehicles and associated systems and methods for transporting crew and cargo to space (e.g., reusable Earth-to-orbit vehicles). 
     BACKGROUND 
     There are a number of existing launch vehicles available for transporting crew and cargo to Low Earth Orbit (LEO) and to existing in-orbit systems, such as the International Space Station (ISS). The Space Shuttle and the SpaceX Falcon 9 are two such vehicles. The Space Shuttle, however, was costly to operate, and although many of its systems were reusable, with the notable exception of the large external tank, they required a great deal of logistics support to refurbish, reassemble, and relaunch. Additionally, both the Space Shuttle and the Falcon 9 were designed to carry relatively heavy payloads of about 50,000 lbs. to LEO. As a result, these vehicles do not present viable, relatively low-cost options for transporting crew and/or lighter cargo (e.g., about 5-10,000 lbs.) to LEO. 
     Space transportation systems include single-stage-to-orbit (SSTO) launch vehicles as well as multi-stage-to-orbit vehicles. In the early 1970&#39;s, Boeing developed a design proposal for a REUSABLE AERODYNAMIC SPACE VEHICLE (RASV). Although the RASV was never built, the proposed design was a SSTO Horizontal Takeoff, Horizontal Landing (HTHL) spaceplane that utilized a sled boost assisted launch. The vehicle was primarily directed toward military space missions, and utilized a welded metal honeycomb airframe with integral thermal protection. The proposed propulsion system included very complex, high maintenance liquid oxygen (LOX)/liquid hydrogen (LH2) Space Shuttle main engines. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG.  1 A  is a partially schematic, top rear isometric view of an aerospace vehicle configured in accordance with embodiments of the present technology, and  FIGS.  1 B- 1 E  are series of front, side, top front isometric, and bottom front isometric views, respectively, of the aerospace vehicle. 
         FIG.  2    is a partially schematic, partially cross-sectioned top view of the aerospace vehicle configured in accordance with embodiments of the present technology. 
         FIG.  3 A  is a cross-sectional side isometric view illustrating a forward portion of the fuselage of the aerospace vehicle, and  FIG.  3 B  is an isometric view of the aerospace vehicle docking with an orbiting station in space, in accordance with embodiments of the present technology. 
         FIGS.  4 A- 4 C  are a series of top rear isometric, top cross-sectional isometric, and side cross-sectional isometric views, respectively, of a rocket powered launch sled configured in accordance with embodiments of the present technology, and  FIG.  4 D  is a schematic diagram illustrating propellant distribution between the launch sled and the vehicle in accordance with an embodiment of the present technology. 
         FIGS.  5 A and  5 B  are partially schematic side and end views, respectively, of a coupling for movably mounting a launch sled to one or more launch rails in accordance with embodiments of the present technology. 
         FIG.  6    is a flow diagram of a method for attaching the aerospace vehicle of  FIGS.  1 A- 1 E  to the launch sled of  FIGS.  4 A- 4 C  in accordance with embodiments of the present technology. 
         FIG.  7 A  is a front isometric view of the aerospace vehicle operably mounted to the launch sled, and  FIGS.  7 B- 7 E  are a series of side views illustrating various stages of a process for attaching the vehicle to the launch sled and preparing the vehicle for launch in accordance with embodiments of the present technology. 
         FIG.  8 A  is a partially schematic side view illustrating forward support arm and electrical umbilical connections to an aerospace vehicle, and  FIG.  8 B  is a partially schematic rear view illustrating rear support arm and propellant umbilical connections to the aerospace vehicle, configured in accordance with embodiments of the present technology. 
         FIGS.  9 A and  9 B  are side and rear views, respectively, of a support arm hold and release mechanism in a first stage of operation, and  FIG.  9 C  is a side view of the support arm hold and release mechanism in a second stage of operation, in accordance with embodiments of the present technology. 
         FIGS.  10 A and  10 B  are a front view and a side cross-sectional view, respectively, of a vehicle support arm interface configured in accordance with embodiments of the present technology. 
         FIGS.  11 A and  11 B  are a front view and a side cross-sectional view, respectively, of a support arm end fitting configured in accordance with embodiments of the present technology. 
         FIGS.  12 A and  12 B  are side and rear views, respectively, of another support arm hold and release mechanism in a first stage of operation, and  FIGS.  12 C and  12 D  are side and rear views, respectively, of the support arm hold and release mechanism in a second stage of operation, in accordance with embodiments of the present technology. 
         FIG.  13    is a schematic diagram of an aerospace vehicle and a launch sled operably coupled to a propellant management system configured in accordance with embodiments of the present technology. 
         FIG.  14    is a flow diagram of a routine for loading propellants onto a launch sled and an aerospace vehicle in accordance with embodiments of the present technology. 
         FIG.  15    is a schematic diagram of a system architecture for controlling a propellant management system, a launch sled, and an aerospace vehicle during propellant loading, in accordance with embodiments of the present technology. 
         FIG.  16 A  is a flow diagram of a routine for operating an aerospace vehicle and a launch sled during a takeoff run;  FIG.  16 B  is a flow diagram of a routine for confirming safe liftoff conditions prior to release of the aerospace vehicle from the launch sled;  FIG.  16 C  is a flow diagram of a routine for liftoff of the aerospace vehicle from the launch sled; and  FIG.  16 D  is a flow diagram of a routine for aborting liftoff of the aerospace vehicle from the launch sled, in accordance with embodiments of the present technology. 
         FIGS.  17 A- 17 D  are a series of schematic diagrams illustrating operation of an aerospace vehicle and a launch sled at various stages of a launch process in accordance with embodiments of the present technology. 
         FIG.  18 A  is a schematic diagram of a system architecture for controlling an aerospace vehicle and a launch sled after separation from a propellant management system; and  FIG.  18 B  is a schematic diagram of a system architecture for controlling the aerospace vehicle after liftoff from the launch sled, in accordance with embodiments of the present technology. 
         FIG.  19    is a partially schematic diagram illustrating various stages of a flight sequence of an aerospace vehicle in accordance with embodiments of the present technology. 
         FIG.  20    is a flow diagram of a routine for performing an ascent of an aerospace vehicle in accordance with embodiments of the present technology. 
         FIG.  21    is a chart listing example types of mission aborts and engine failures/degradations that an aerospace vehicle could experience, in accordance with embodiments of the present technology. 
         FIG.  22    is a flow diagram of a routine for responding to an engine anomaly after liftoff of an aerospace vehicle in accordance with embodiments of the present technology. 
         FIG.  23 A  is a schematic diagram of a control system for an aerospace vehicle, and  FIG.  23 B  is a schematic diagram of a control system for a launch sled, configured in accordance with embodiments of the present technology. 
         FIG.  24 A  is a partially schematic diagram of an oxidizer tank configured in accordance with embodiments of the present technology, and  FIG.  24 B  presents a graph of various pressures versus time for the oxidizer tank and its environment. 
         FIG.  25 A  is a partially schematic diagram of a fuel tank configured in accordance with embodiments of the present technology, and  FIG.  25 B  presents a graph of various pressures versus time for the fuel tank and its environment. 
         FIGS.  26 A and  26 B  are top front isometric and bottom front isometric views, respectively, of the aerospace vehicle of  FIGS.  1 A- 1 D  illustrating various aspects of the airframe and an associated thermal protection system, configured in accordance with embodiments of the present technology. 
     
    
    
     DETAILED DESCRIPTION 
     The following disclosure describes various embodiments of space launch vehicle systems and associated methods of manufacture and use. In some embodiments, the systems include a fully reusable, horizontal takeoff/horizontal landing (HTHL), ground-assisted single-stage-to-orbit (SSTO) spaceplane that is capable of providing frequent deliveries of people and/or cargo to Low Earth Orbit (LEO). As described in greater detail below, the spaceplane can take off with the aid of a rocket-powered sled that, in addition to providing thrust for takeoff, can also provide propellant for the spaceplane engines during the takeoff run so that the spaceplane launches with full propellant tanks. In some embodiments, the sled can utilize magnetic levitation and magnetic propulsion to provide thrust for takeoff. After several hours or days in orbit, the spaceplane can fly back to Earth and land on a conventional runway having a length of, for example, about 10,000 ft. Embodiments of the systems disclosed herein can enable the expansion of the existing space industry by providing low-cost access to orbital destinations, such as the International Space Station (ISS), for people and light to medium cargo (e.g., about 5,000 lbs.) at relatively high frequency (e.g., as often as twice a week). 
     Certain details are set forth in the following description and in  FIGS.  1 - 26 B  to provide a thorough understanding of various embodiments of the present technology. In other instances, well-known structures, materials, operations and/or systems often associated with aerospace vehicle structures, propulsion systems, control systems, flight sequences, control routines, etc. are not shown or described in detail in the following disclosure to avoid unnecessarily obscuring the description of the various embodiments of the technology. Those of ordinary skill in the art will recognize, however, that the present technology can be practiced without one or more of the details set forth herein, or with other structures, methods, components, and so forth. 
     The accompanying Figures depict embodiments of the present technology and, unless otherwise specified, are not intended to be limiting of its scope. The sizes of various depicted elements are not necessarily drawn to scale, and these various elements may be arbitrarily enlarged to improve legibility. Component details may be abstracted in the Figures to exclude details such as position of components and certain precise connections between such components when such details are unnecessary for a complete understanding of how to make and use the disclosed technology. Many of the details, dimensions, angles and other features shown in the Figures are merely illustrative of particular embodiments of the disclosure. Accordingly, other embodiments can have other details, dimensions, angles and features without departing from the present technology. In addition, those of ordinary skill in the art will appreciate that further embodiments of the present technology can be practiced without several of the details described below. 
     In general, identical reference numbers in the Figures identify identical, or at least generally similar, elements. To facilitate the discussion of any particular element, the most significant digit or digits of any reference number refers to the Figure in which that element is first introduced. For example, element  110  is first introduced and discussed with reference to  FIG.  1   . 
       FIG.  1 A  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.  FIGS.  1 B- 1 E  are front, side, top front isometric, and bottom front isometric views, respectively, of the vehicle  100 . Referring first to  FIGS.  1 A- 1 C , in the illustrated embodiment the vehicle  100  is an HTHL/SSTO 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  110  (identified individually as a left vertical stabilizer  110   a  and a right vertical stabilizer  110   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  103  in an upper portion of a crew cabin  112  for crew ingress and egress. Additionally, in the illustrated embodiment the forward portion of the fuselage  102  includes a movable hatch  114  for providing access to a docking port (not shown in  FIG.  1 A- 1 E ) for docking the vehicle  100  with an on-orbit station, such as the ISS, and enabling human and/or cargo movement therebetween. 
     As shown in  FIG.  1 A , in the illustrated embodiment the aft portion of the fuselage  102  carries a propulsion system  111  having one or more rocket engines  120  (three are shown in  FIG.  1 A ; identified individually as first, second, and third rocket engines  120   a - c , respectively). Each of the engines  120   a - c  has a corresponding nozzle  117  positioned generally proximate the trailing edge portion of the wings  104   a, b  between the vertical stabilizers  110   a, b . As described in greater detail below, in some embodiments the rocket engines  120   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  120   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  117 , 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  117  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. In some embodiments, the engines  120   a - c  can be at least generally similar in structure and function to engines described in U.S. Provisional Application No. 62/693,829, filed on Jul. 3, 2018, and titled “ROCKET PROPULSION SYSTEMS AND ASSOCIATED METHODS,” which is incorporated herein by reference in its entirety. 
     The vehicle  100  can further include orbital maneuvering system (OMS) engines  122  (identified individually as a first OMS engine  122   a  and a second OMS engine  122   b ) having nozzles positioned just above the nozzles for the main engines  120   a - c . In some embodiments, the OMS engines  122  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  122 . As described in greater detail below, 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  120  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  140  having one or more processors  142  that can control various operations and functions of the vehicle  100  in accordance with computer-readable instructions stored on system memory  144 . The controller  140  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  142  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  142  may be a single processing unit or multiple processing units distributed across multiple systems and/or subsystems of the vehicle  100 . The processor  142  is operably connected to the memory  144  and may be operably connected to various systems of the vehicle  100  to transmit instructions and/or receive input therefrom. The memory  144  can include read-only memory (ROM) and random-access memory (RAM) 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  146  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  146  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  148 ; a navigation system  150 ; a flight controls system  152  for affecting actuation of the vehicle control surfaces, engine throttles, landing gear, etc.; avionics  154 ; a hydraulic system  156  for, e.g., control surface and landing gear actuation; and an environmental control system  158  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. 
     As shown in  FIGS.  1 B and  1 C , in the illustrated embodiment the vehicle  100  includes a landing gear system having a nose gear  126   a , a left main gear  126   b , and a right main gear  126   c . (In  FIG.  1 E , the landing gear  126   a - c  are retracted and stowed behind corresponding gear doors  129   a - c .) As described in more detail below, the vehicle  100  takes off with the assistance of a launch sled and, as a result, the landing gear  126   a - c  are retracted and stowed into associated gear bays during takeoff. For landing, the landing gear  126   a - c  are deployed in a manner that is at least generally similar to conventional commercial aircraft. Since the landing gear  126   a - c  are only designed to carry the loads associated with landing the vehicle  100  when it is not carrying a full load of propellant and is therefore relatively light, the landing gear  126   a - c  can be substantially lighter than they would otherwise be if they were designed to support the vehicle  100  during takeoff with a full load of propellant. This weight savings results in an increased payload capacity. 
     Referring next to  FIG.  1 D , in some embodiments the vehicle  100  can include an antenna  134  which can be deployed from the fuselage  102  once the vehicle  100  is in orbit to facilitate communications between Earth, on-orbit stations, and/or other remote locations. The vehicle  100  can further include a plurality of thrusters  136  positioned at various locations on the exterior of the vehicle  100  to provide attitude control while on orbit. Such reaction control system (RCS) thrusters can include, for example, relatively small monopropellant thrusters known in the art. In some embodiments, the vehicle  100  may be configured to use only “green” propellants. In such embodiments, the thrusters  136  can use a hydroxylammonium nitrate-based propellant known as AF-M315E, and/or a propellant known as LMP-103S, which is based on the oxidizer ammonium dinitramide. Both of these propellants are less toxic than, for example, hydrazine. In other embodiments, the thrusters  136  can use hydrazine in a conventional manner. In some embodiments, the vehicle  100  can include a pair of thrusters  136   a, b  toward an upper aft portion of the fuselage  102 , and a pair of thrusters  136   c, d  near the tip portion of each of the wings  104   a, b . Additionally, the vehicle  100  can also include a pair of thrusters  136   e, f  at the base of each vertical stabilizer  110   a, b , a pair of thrusters  136   g, h  toward an upper forward portion of the fuselage  102  just behind the crew cabin  112 , and another pair of thrusters  136   i, j  on opposite sides of the fuselage near the same location. As shown in  FIG.  1 E , additional thrusters  136   k, l  and  136   m, n  can also be included on the underside of the vehicle  100  proximate the forward and aft portions the fuselage  102 , respectively. Selective activation of the various thrusters  136  enable the vehicle to be positioned in virtually any attitude while in orbit to facilitate, for example, docking with on-orbit stations, transmission and/or reception of communications, planetary viewing, etc. The thruster locations illustrated in  FIGS.  1 D and  1 E  are provided by way of examples of some embodiments. Accordingly, it will be understood that other embodiments can have more or fewer thrusters and/or thrusters in other locations. 
     As also shown in  FIG.  1 E , in the illustrated embodiment the vehicle  100  further includes propellant interfaces  124  (identified individually as a first propellant inlet interface  124   a  and a second propellant inlet interface  124   b ), and support arm interfaces  127  (identified individually as a forward support arm interface  127   a , a first aft support arm interface  127   b , and a second aft support arm interface  127   c ). Each of the support arm interfaces  127   a - c  can include a coupling configured to releasably engage a corresponding support arm for mounting the vehicle  100  to the launch sled (not shown in  FIG.  1 E ) prior to and during takeoff. In some embodiments, each of the propellant inlet interfaces  124   a, b  is positioned just inboard of (and laterally adjacent to) the corresponding support arm interface  127   b, c . As described in greater detail below, each of the propellant interfaces  124   a, b  can include a quick-disconnect valve and/or other suitable coupling for releasably connecting a corresponding propellant umbilical (e.g., a propellant conduit; which can also be referred to as a propellant line) from the launch sled to the vehicle  100  and sealing the interface when the propellant umbilical is disconnected. Additionally, the propellant inlet interfaces  124   a, b  can also include doors that close flush with the outer surface of the fuselage  102  to protect the interfaces  124   a, b  from aerothermal heating, etc. Propellants (e.g., LOX and Jet-A) are transferred from the launch sled to the vehicle  100  via the propellant umbilicals and the inlet interfaces  124   a, b  for operation of the main vehicle engines  120   a - c  during takeoff. For example, in some embodiments the first propellant inlet interface  124   a  can be configured to receive LOX from the launch sled via a LOX umbilical, and the second propellant inlet interface  124   b  can be configured to receive fuel from the launch sled via a fuel umbilical. Additionally, as described in greater detail below, in some embodiments the first propellant inlet interface  124   a  can also be configured to recirculate vented/warmed LOX from the vehicle  100  back to the launch sled via the LOX umbilical, and the second propellant inlet interface  124   b  can also be configured to recirculate vented fuel from the vehicle  100  back to the launch sled via the fuel umbilical. 
     In addition to the propellant inlet interfaces  124   a, b , the vehicle  100  further includes an electrical interface  125  positioned just aft of the forward support arm interface  127   a . The electrical interface  125  is configured to releasably connect to an electrical umbilical that extends from the launch sled to the vehicle  100 . As described in greater detail below, the electrical interface  125  can include one or more electrical receptacles configured to receive one or more corresponding connectors on the electrical umbilical to enable transmission of commands, power, and data between the vehicle  100  and the launch sled. Like the propellant interfaces  124   a, b , the electrical interface  125  can also include a door that closes flush with the outer surface of the fuselage  102  after the electrical umbilical has been disconnected to protect the interface  125  from aerothermal heating, etc. 
       FIG.  2    is a partially schematic, partially cross-sectioned top view of the vehicle  100  configured in accordance with embodiments of the present technology. In one aspect of this embodiment, the oxidizer (e.g., LOX) for the main engines  120   a - c  is contained in a fuselage tank  242  that is formed by external sidewalls of the fuselage  102  and, accordingly, the tank  242  can have a cross-sectional shape that follows the cross-sectional shape (contour) of the fuselage  102 . More specifically, as discussed in greater detail below, in some embodiments the LOX is subcooled to a temperature of, for example, about −320 degrees Fahrenheit (F) (i.e., about −196 degrees Celsius (C)). At this temperature, the LOX vapor pressure is sufficiently low that the pressure differential across the walls of the fuselage tank  242  is less than about 3 psig, such as between 2-3 psig. By maintaining the tank pressure at about 2-3 psig, the structural loads on the tank walls are relatively low. As a result, the tank  242  does not have to have the shape of a conventional high-pressure propellant tank or pressure vessel (e.g., a spherical shape or a cylindrical shape having a circular cross-section). This enables the internal volumes of the airframe (e.g., the fuselage  102 ) to be used as a LOX tank, while at the same time being shaped for optimum aerodynamic performance without requiring any structural reinforcement to accommodate high tank pressure loads. For example, in some embodiments the oxidizer tank  242  can have a cross-sectional shape that is non-circular, such as an oval, or near oval, cross-sectional shape, an elliptical cross-sectional shape, an asymmetric cross-sectional shape, and/or other non-circular cross-sectional shapes. Additionally, as shown by the plan view of  FIG.  2   , in some embodiments the cross-section of the oxidizer tank  242  can vary in both shape and/or size along the length of the fuselage  102 . In other embodiments, however, the oxidizer tank  242  can have other shapes, such as cylindrical and/or spherical shapes. The oxidizer tank  242  is connected in fluid communication with the first propellant inlet interface  124   a  ( FIG.  1 E ). In another aspect of this embodiment, each of the wings  104   a, b  includes a corresponding fuel tank  240   a, b  that contains the fuel for the vehicle main engines  120   a - c . The fuel tanks  240   a, b  fill much of the interior volumes of the wings  104   a, b  in the forward strake regions and the main wing regions, except for the volumes proximate wing leading edge regions  246  and main spar sections  248 . The fuel tanks  240   a, b  are connected in fluid communication with the second propellant inlet interface  124   b  ( FIG.  1 E ). The OMS propellant tanks  244  (e.g., LOX and CNG tanks) are positioned between the main engine oxidizer tank  242  and the main engines  120   a - c.    
       FIG.  3 A  is a side cross-sectional view of the forward portion of the fuselage  102  illustrating the crew cabin  112  and an adjacent payload bay  374 . In some embodiments, the payload bay  374  can include an airlock and a docking port  358 . In other embodiments, in addition to or instead of the docking port  358 , the payload bay  374  can include a payload support and deployment system (not shown) configured to carry payloads (e.g., one or more satellites) and deploy them into orbit.  FIG.  3 B  is an isometric view of an orbiting station  368  operably coupled to the vehicle  100  via the docking port  358  in accordance with embodiments of the present technology. Referring first to  FIG.  3 A , in the illustrated embodiment the crew cabin  112  can include a plurality of seats  350  (e.g.,  5  seats) for vehicle crew and/or passengers. The crew cabin  112  can also include a plurality of windows  370  for occupant viewing outside of the vehicle  100 . 
     In some embodiments, the crew cabin  112  is a self-contained unit that can separate from the rest of the fuselage  102  in the event of a mission critical failure that occurs at any point during flight. More specifically, an aft portion of that crew cabin  112  can be sealably enclosed by a pressure bulkhead  354  that enables the crew cabin  112  to maintain internal pressure during all phases of operation. Additionally, the crew cabin  112  can be structurally attached to the rest of the fuselage  102  by a frangible joint  356  that extends around the circumference of the fuselage just aft of the bulkhead  354 . The frangible joint  356  can include a pyrotechnically actuated explosive device (e.g., Super Zip from Ensign-Bickford Aerospace Company) or linear shaped charge that structurally attaches the crew cabin  112  to the fuselage  102  until actuated in response to a separation signal. Upon actuation, the frangible joint  356  breaks to immediately detach the crew cabin  112  from the rest of the fuselage  102 . In other embodiments, instead of (or in addition to) the frangible joint  356 , the crew cabin  112  can be attached to the rest of the fuselage  102  with a plurality of explosive bolts and/or other known separating devices to enable the crew cabin  112  to be quickly disengaged and separated from the fuselage  102  in the event of a mission critical failure of one or more vehicle systems. 
     The vehicle  100  can include a number of subsystems to facilitate separation of the crew cabin  112  from the rest of the fuselage  102  and safe return of the crew cabin  112  to Earth in the event of a mission critical failure. For example, in the illustrated embodiment the crew cabin  112  can include a recovery chute  364  positioned toward an upper aft portion of the crew cabin  112 , and a downward-firing thruster  360   a  and a first aft-firing thruster  360   b  positioned toward a lower aft portion of the crew cabin  112 . Although not shown by virtue of the section view, a second aft-firing thruster is positioned adjacent to the first aft-firing thruster  360   b  on the opposite sided of the vehicle centerline. The thrusters  360   a, b  can be conventional bipropellant or green propellant thrusters that receive propellant from corresponding fuel and oxidizer tanks  362  positioned beneath a floor  372  of the crew cabin  112 . The nozzle of the first thruster  360   a  can be directed generally downward and aft, and the nozzle of the second thruster  360   b  can be directed generally aft. The recovery chute  364  can include one or more parachutes that are deployed from the crew cabin  112  after separation from the rest of the fuselage  102  and when the crew cabin  112  is at an appropriate altitude on descent. 
     As described in greater detail below, in the event the vehicle  100  experiences a mission critical failure at any point during flight, the frangible joint  356  can be activated to separate the crew cabin  112  from the rest of the fuselage  102 . Immediately after separation, the thrusters  360   a, b  can be ignited to quickly move the crew cabin  112  a safe distance away from the rest of the vehicle  100 . Once the crew cabin  112  descends to an appropriate altitude, the recovery chute  364  can be deployed to slow the decent through the Earth&#39;s atmosphere. The recovery chute  364  can be positioned to properly orient the crew cabin  112  during descent through the atmosphere. Additionally, the crew cabin  112  can also include a thermal protection system (e.g., a fibrous, reinforced oxidation-resistant composite covering) on at least the forward-facing surfaces to provide sufficient heat shielding during reentry. 
     In another aspect of this embodiment, the crew cabin includes a passage  352  (with a sealable door, not shown) that enables the crew of the vehicle  100  to move back and forth between the crew cabin  112  and the adjacent airlock or payload bay  374  when in orbit. As shown in  FIG.  3 B , in operation the hatch  114  can be opened to enable the docking port  358  to sealably engage and structurally attach to a corresponding docking port  366  on the orbiting station  368 . Once properly docked, crew and/or cargo from the vehicle  100  can move back and forth between the vehicle  100  and the orbiting station  368 . 
       FIG.  4 A  is a partially schematic isometric view of a launch sled  400  configured in accordance with embodiments of the present technology, and  FIGS.  4 B and  4 C  are top cross-sectional isometric and side cross-sectional isometric views, respectively, of the launch sled  400 . Referring first to  FIG.  4 A , the sled  400  runs on three heavy duty rails  410   a - c  that react the combined loads from the sled  400  and the vehicle  100  during acceleration and deceleration. In some embodiments, the rails  410   a - c  can be about two miles long. In the illustrated embodiment, the sled  400  includes a body or chassis  402  having a center section  446   a , a left outer section  446   b , and a right outer section  446   c  that are connected by an aerodynamic top plate  406 . The sections  446   a - c  can also be referred to as “trucks.” The underside of each of the sections  446   a - c  is moveably engaged with a corresponding rail  410   a - c  by at least a forward coupling  454   a  and an aft coupling  454   b . In operation, the couplings  454   a, b  enable the sled to move fore and aft on the rails  410   a - c , while keeping the sled  400  attached to the rails. 
     The sled  400  further includes a forward support arm  416   a  operably coupled to a forward portion of the center section  446   a , and first and second aft support arms  416   b  and  416   c , respectively, operably coupled to the first outer section  446   b  and the second outer section  446   c , respectively. In some embodiments, the proximal end portion of each of the support arms  416   a - c  is pivotably coupled to the sled chassis  402 , and the arms  416   a - c  are operably coupled to a drive system (e.g., an electromechanical system, a hydraulic system, a pneumatic system, etc.; not shown) that moves (e.g., rotates) the arms  416   a - c  through their ranges of operating motion. The distal end portion of each support arm  416   a - c  is configured to be releasably attached to the corresponding support arm interface  127   a - c  on the vehicle  100  by means of a hold and release mechanism. For example, in some embodiments the distal end portion of each support arm  416   a - c  can include a fitting  422  (e.g., a ball fitting or other suitable fitting that can permit rotation of the arm  416  while the vehicle  100  is attached) that is configured to be releasably engaged with the corresponding support arm interface  127   a - c  (e.g., a ball socket) by the hold and release mechanism. In some embodiments, the hold and release mechanism can include a mechanical clamp mechanism that holds the distal end portion of each control arm  416   a - c  to the corresponding interface  127   a - c  until commanded to release. In other embodiments, other suitable hold and release mechanisms known in the art can be used to releasably attach the support arms  416   a - c  to the corresponding interfaces  127   a - c.    
     In  FIG.  4 A , each of the support arms  416   a - c  is illustrated in three different operational positions: a lowermost or stowed position, an intermediate position in which the arm is rotated partially upward to engage and raise the vehicle  100  for stowage of the landing gear  126   a - c , and a launch position in with the arm is rotated fully aft and/or upward to optimize (or nearly optimize) vehicle angle of attack for separation and lift off. The support arms  416   a - c  react all the acceleration and deceleration loads between the vehicle  100  and the sled  400  during takeoff, thereby eliminating the need for the vehicle  100  to include a landing gear system that is rated for takeoff loads. Additionally, the positioning of the support arms  416   a - c  can be adjustable to optimize the vehicle angle of attack as needed for each mission. Although the illustrated embodiment includes three support arms  416   a - c , in other embodiments the sled  400  can include more support arms (e.g., four, five, or more arms) or fewer arms (e.g., two or one arm). 
     In the illustrated embodiment, the sled  400  includes three rocket engines  404  (identified individually as rocket engines  404   a - c ) which are mounted to the aft portions of the corresponding sled sections  446   a - c . In other embodiments, the sled  400  can have more or fewer rocket engines (e.g., one to five or more rocket engines). The rocket engines  404   a - c  have nozzles with area ratios optimized for ground performance. Like the main vehicle engines  120   a - c  ( FIG.  1 A ), the engines  404   a - c  can be bipropellant rocket engines configured to burn, for example, LOX and Jet-A. Since the engines  404   a - c  operate at a constant or near-constant altitude, their nozzles can be configured for single area ratio operation and optimum performance at the altitude at which the sled operates (e.g., sea level or near sea level). In some embodiments, the engines  404   a - c  can be at least generally similar in structure and function to engines described in U.S. Provisional Application No. 62/693,829, filed on Jul. 3, 2018, and titled “ROCKET PROPULSION SYSTEMS AND ASSOCIATED METHODS,” which is incorporated herein by reference in its entirety. 
     In some embodiments, each of the engines  404   a - c  can have dedicated propellant tanks (e.g., a dedicated oxidizer (e.g., LOX) tank  452  and/or a dedicated fuel (e.g., Jet-A) tank  450 ) mounted and housed within the enclosure of the corresponding sled section  446   a - c , as shown in  FIGS.  4 B and  4 C . The propellant tanks  450  and  452  are sized and configured to provide propellant to the corresponding engines  404   a - c  for the duration of the sled run. In other embodiments, the sled  400  can carry a single oxidizer tank and/or a single fuel tank that provides sufficient propellant for all the engines  404   a - c . In further embodiments, all or a portion of the engines  404   a - c  may use different propellants, and/or one or more of the engines  404   a - c  may be solid rocket motors. 
     In addition to the sled propellant tanks  450  and  452 , the sled  400  can also carry auxiliary propellant tanks (e.g., an oxidizer (e.g., LOX) tank  458  and a fuel (e.g., Jet-A) tank  456 ; shown in cross-section in  FIGS.  4 B and  4 C ) for providing propellant to the vehicle engines  120   a - c  ( FIG.  1 A ) for the duration of the sled run. As described in greater detail below, this assures that the vehicle  100  is as full of propellant as possible at liftoff from the sled, thereby minimizing (or at least greatly reducing) the vehicle dry mass penalty at lift off. In some embodiments, the auxiliary propellant tanks  456  and  458  can be housed in the center section  446   a . In other embodiments, the propellant tanks  456  and  458  can be carried in one or both of the left outer section  446   b  and/or the right outer section  446   c . In other embodiments, one or both of the auxiliary propellant tanks  456  and  458  can be omitted, and one or more of the sled propellant tanks  452  and  450  can be sized and configured to provide propellant to both the corresponding sled engine(s)  404  as well as the vehicle engines  120   a - c  for the duration of the sled run. 
     The sled  400  further includes a first propellant outlet interface  414   a  and a second propellant outlet interface  414   b  located on or proximate the top plate  406 . In the illustrated embodiment, the first propellant outlet interface  414   a  is positioned inboard of, and laterally adjacent to, the base of the first aft support arm  416   b  and is operably connected in fluid communication with the auxiliary oxidizer tank  458 . The second propellant outlet interface  414   b  is positioned inboard of, and laterally adjacent to, the base of the second aft support arm  416   c  and is operably connected in fluid communication with the auxiliary fuel tank  456 . A first propellant umbilical  460   a  (e.g., a LOX umbilical) extends from the first propellant outlet interface  414   a  and has an outlet  461   a , and a second propellant umbilical  460   b  (e.g., a fuel umbilical) extends from the second propellant outlet interface  414   b  and has an outlet  461   b . As described above with reference to  FIG.  1 E , in operation the outlet  461   a  of the first propellant umbilical  460   a  is configured to releasably connect to the first propellant inlet interface  124   a  on the vehicle  100 , and the outlet  461   b  of the second propellant umbilical  460   b  is configured to releasably connect to the second propellant inlet interface  124   b  on the vehicle  100 . 
     As described in greater detail below, during launch of the vehicle  100 , the auxiliary oxidizer tank  458  provides oxidizer (e.g., LOX) to the vehicle oxidizer tank  242  ( FIG.  2   ) for the vehicle main engines  120   a - c  via the first propellant umbilical  460   a , and the auxiliary fuel tank  456  provides fuel (e.g., Jet-A) to the vehicle fuel tanks  240   a, b  ( FIG.  2   ) for the vehicle main engines  120   a - c  via the second propellant umbilical  460   b . Additionally, as further described below, in some embodiments the first propellant umbilical  460   a  is also configured to recirculate vented oxidizer from the vehicle oxidizer tank  242  back to the sled  400 , and the second propellant umbilical  460   b  is also configured to recirculate vented fuel from the vehicle fuel tanks  240   a, b  back to the sled  400 . Just prior to separation of the vehicle  100  from the sled  400 , the valves associated with each of the vehicle propellant interfaces  124   a, b  ( FIG.  1 E ) are closed, the propellant umbilicals  460   a, b  from the sled  400  are disconnected from the corresponding interfaces  124   a, b , and the associated interface doors are closed. By launching the vehicle  100  in this way, the sled  400  operates as both a “first stage” and an external propellant tank of the vehicle  100 . 
     In addition to the propellant interfaces  414   a, b , the sled  400  can further include an electrical umbilical  426  extending from an electrical interface  425  positioned just aft of the base of the forward support arm  416   a  on the top plate  406 . As described in greater detail below, the distal end portion of the electrical umbilical  426  can include one or more electrical connectors  427  configured to releasably connect to corresponding receptacles on the vehicle electrical interface  125  ( FIG.  1 E ) so that operating commands, power and data can be transmitted between the sled  400  (e.g., the sled controller  440 ) and the vehicle  100  while the vehicle is mounted to the sled. 
     The sled  400  can further include a sled braking system  455  configured to slow the sled  400  to a stop if an anomalous operating condition of one or more of the sled engines  404   a - c  or the main vehicle engines  120   a - c  is detected during launch. The sled braking system  455  also slows the sled to a stop at the end of each takeoff run. In some embodiments, the braking system  455  can be at least generally similar and structure and function to the braking system used on the Holloman high speed test track at Holloman Air Force Base in New Mexico, U.S.A., which includes a water brake comprising a series of water barriers of calibrated depth that slow and stop the sled. In other embodiments, the braking system  455  can include hydraulically-actuated brakes (not shown) that engage the rails  410   a - c  to slow the sled. In other embodiments, the sled  400  can include other mechanically, pneumatically, electrically, magnetically, and/or hydraulically actuated forms of braking devices including, for example, reverse thrust rocket engines (also not shown). 
     The sled  400  can also include a controller  440  having a processor  442  and memory  444 . The controller  440  can receive inputs  413  and issue outputs  415  (e.g., commands directing sled operation, including operating valves, pumps, actuators, and/or other components). By way of example, the inputs  413  can include control signals and commands from, e.g., ground systems, the vehicle  100 , etc., data from the vehicle  100 , operating parameters such as speed, temperature, etc., engine operating parameters, propellant parameters, and/or other information). The outputs  415  can include commands directing sled operation (including engine operation, e.g., start, stop and throttle settings, braking system operation, etc.), data transmissions, etc. The processor  442  can include one or more logic processing units, such as one or more CPUs, DSPs, ASICs, etc., that are operably connected to controls associated with, for example, the engines  404   a - c , the propellant tanks  450 ,  452 ,  456  and  458 , the support struts  416   a - c , the sled braking system  455 , etc. The processor  442  can control operation of these sled systems as described herein in accordance with computer-readable instructions stored on the memory  444 . 
       FIG.  4 D  is a schematic diagram illustrating propellant distribution between the launch sled  400  and the vehicle  100  during launch, in accordance with an embodiment of the present technology. In this particular embodiment, one or more of the sled oxidizer tanks  452  and the sled fuel tanks  450  provide propellant for both the vehicle engines  120   a - c  and the corresponding sled engine(s)  404   a - c  during the sled run. In  FIG.  4 D , the vehicle  100  is mounted to the launch sled  400  in preparation for launch. For ease of illustration, the oxidizer tanks  452  and the fuel tanks  450  on the launch sled  400  are represented by a single oxidizer tank  452  and a single fuel tank  450 , respectively. In the illustrated embodiment, the launch sled  400  includes a first propellant conduit or line  462   a  that provides fuel from the fuel tank  450  to the sled engines  404   a - c , and a second propellant line  462   b  that provides oxidizer from the oxidizer tank  452  to the engines  404   a - c . Although not shown, one or more propellant pumps can be associated with each of the tanks  450 ,  452  to drive propellant from the tanks through the associated propellant lines  462   a, b , and to the vehicle  100 . Additionally, the launch sled  400  further includes a first valve  464   a  in fluid communication with the first line  462   a  that can direct fuel from the fuel tank  450  to the first propellant outlet interface  414   a , and a second valve  464   b  in fluid communication with the second line  462   b  that can direct oxidizer from the oxidizer tank  452  to the second propellant outlet interface  414   b . Prior to launch, the first propellant line  460   a  is operably connected to the first propellant inlet interface  124   a  on the vehicle  100 , and the second propellant line  460   b  is operably connected to the second propellant inlet interface  124   b.    
     In the illustrated embodiment, the vehicle  100  includes a first propellant line  466   a  that provides oxidizer from the vehicle fuel tanks  240   a, b  to the vehicle main engines  120   a - c , and a second propellant line  466   b  that provides fuel from the vehicle oxidizer tank  242  to the main engines  120   a - c . Although not shown, one or more propellant pumps can be associated with each of the tanks  240   a, b  and  242  to drive propellant from the tanks through the associated propellant lines  466   a, b . Additionally, the vehicle  100  further includes a first valve  468   a  in fluid communication with the first propellant line  466   a , and a second valve  468   b  in fluid communication with the second propellant line  466   b . The first valve  468   a  is configured to receive fuel from the first propellant inlet interface  124   a  and provide the fuel to the engines  120   a - c . Similarly, the second valve  468   b  is configured to receive oxidizer from the second propellant inlet interface  124   b  and provide the oxidizer to the engines  120   a - c.    
     Propellant can be distributed between the launch sled  400  and the vehicle  100  during vehicle takeoff in one embodiment as follows. Prior to takeoff, the valves  464   a, b  are opened to provide fuel and oxidizer from the fuel and oxidizer tanks  450  and  452 , respectively, to the sled engines  404   a - c  for ignition. Additionally, the valves  464   a, b  also provide fuel and oxidizer to the valves  468   a, b , respectively, on the vehicle  100  via the propellant lines  460   a, b , respectively. Prior to igniting the sled engines  404   a - c , the valves  468   a, b  are positioned to direct the fuel and oxidizer from the sled fuel tank  450  and the sled oxidizer tank  452 , respectively, to the vehicle main engines  120   a - c  for ignition. The sled engines  404   a - c  and the vehicle engines  120   a - c  are then ignited, and once the sled and vehicle engines come up to launch thrust, the sled  400  is released on the takeoff run down the rails  410   a - c  ( FIG.  4 A ). Accordingly, the sled fuel tank  450  and sled oxidizer tank  452  provide fuel and oxidizer, respectively, to both the sled engines  404   a - c  and the vehicle main engines  120   a - c  throughout ignition and the takeoff run of the vehicle  100 . Just prior to (or during) separation of the vehicle  100  from the launch sled  400  at the end of the takeoff run, the valves  464   a, b  on the sled  400  are closed, and the propellant lines  460   a, b  are disconnected from the corresponding propellant inlet interfaces  124   a, b . Just prior to this time, however, the valves  468   a, b  on the vehicle  100  are positioned to close off the connections to the propellant inlet interfaces  124   a, b  and instead enable propellant to flow from the vehicle oxidizer tank  242  and the vehicle fuel tanks  240   a, b  to the main engines  120   a - c  via the first line  466   a  and the second line  466   b . Thus, the vehicle  100  does not begin burning its own propellant until just prior to separation from the launch sled  400 .  FIG.  4 D  illustrates one approach to propellant distribution between the launch sled  400  and the vehicle  100 , and in other embodiments other approaches can be employed, as described in more detail below. 
       FIGS.  5 A and  5 B  are side and end views, respectably, of the coupling  454   a  configured in accordance with embodiments of the present technology. For ease of reference, the coupling  454   a  will be referred to herein as the “coupling  454 ,” with the understanding that the illustrated embodiment can apply to both of the couplings  454   a  and  454   b  shown in  FIG.  4 A . In the illustrated embodiment, the coupling  454  includes a body  564  and a plurality of sacrificial inserts  560   a - c . The body  564  includes a lug portion  569  that extends upwardly from a pair of legs  570   a  and  570   b . The body  564  is structurally coupled to the underside of the sled  400  by means of a pin or bolt that extends through an opening  566  in the lug portion  569 . Each of the legs  570   a, b  includes a plurality of leg portions  568   a - c  that wrap around an upper cap  562  of the launch rail  410  to couple the coupling  454  to the rail  410 . In one aspect of this embodiment, each of the sacrificial inserts  560   a - c  can be manufactured from a suitably tough synthetic resin, such as polytetrafluoroethylene (PTFE). One example of such material is commonly referred to as Teflon™. Each of the leg portions  568   a - c  has a corresponding sacrificial insert  560   a - c  fixedly attached thereto to provide low friction, load-bearing contact surfaces between the body  564  and the launch rail  410 . As noted above with reference to  FIG.  4 A , the coupling  454  moveably (e.g., slideably) engages the sled  400  with the launch rail  410 , and enables the sled  400  to move fore and aft on the rail  410  while restricting movement perpendicular to the rail (e.g., side-to-side and up-and-down). In operation, the sacrificial inserts  560   a - c  are able to withstand the high temperatures and high pressures that result from supporting the weight of the vehicle  100  and the sled  400  on the launch rails  410  during the vehicle takeoff run. After each use, the sacrificial inserts  560   a - c  can be inspected to determine condition and easily replaced if necessary. 
     In other embodiments, the launch sled  400  can be moveably attached to the launch rails  410   a - c  ( FIG.  4 A ) using other types of suitable coupling devices. For example, in some embodiments the launch sled  400  can be movably coupled to the rails  410   a - c  by a magnetic levitation (“Maglev”) system that is at least generally similar in structure and function to the Maglev systems found on high speed trains. Such Maglev systems are described in, for example, U.S. Pat. No. 6,044,770, titled “Integrated High Speed MAGLEV System,” which is incorporated herein by reference in its entirety. The Maglev system can be incorporated into the rails  410   a - c  and can include, for example, a plurality of magnets to support the sled  400  at relatively high speeds with relatively little, if any, friction between the sled  400  and the rails  410   a - c . Additionally, the Maglev system can also include a plurality of magnets for propelling the sled  400  down the rails  410   a - c  during launch, and/or decelerating and stopping the sled  400  at the end of the launch run. 
     In some embodiments, use of magnetic levitation and/or propulsion can supplement or replace the propulsion provided by the rocket engines  404   a - c  ( FIG.  4 A ). For example, an electromagnetically levitated rocket propelled sled  400  can reduce the amount of propellant carried on the sled that is required to reach takeoff velocity and can also reduce maintenance required for the sled system. In some embodiments, the sled  400  can include a super conducting electromagnetically levitated (SCMaglev) and propelled sled that will eliminate the need for rocket engines on the sled  400 . Such embodiments can significantly reduce the total amount of propellant required for takeoff, reduce launch noise, and improve launch turnaround time so that multiple launches can be carried out in a short period of time (e.g., in a single day). In yet other embodiments, the sled  400  can be moveably coupled to the rails  410   a - c  using a system of rollers. However, if rollers or similar systems are used, the rollers must be carefully selected to ensure that the bearings do not overheat and fail at the sled speeds required for vehicle take off run. 
       FIG.  6    is a flow diagram of a method  600  for mounting the vehicle  100  to the launch sled  400  and preparing the vehicle  100  for launch, in accordance with embodiments of the present technology.  FIG.  7 A  is a front isometric view of the vehicle  100  mounted to the launch sled  400 , and  FIGS.  7 B- 7 E  are a series of side views illustrating various stages of the method of  FIG.  6    in accordance with embodiments of the present technology. Referring to  FIG.  6   , the method  600  begins in block  602  with the vehicle  100  being towed onto the sled  400  on its landing gear  126   a - c , as shown in  FIG.  7 B . In block  604 , command and telemetry umbilicals are connected from the sled  400  to the vehicle  100  and verified, and in block  606 , electrical umbilicals are connected from the sled  400  to the vehicle  100  to provide ground power. In some embodiments, the operations of blocks  604  and  606  can be performed concurrently by a ground crew by connecting the electrical umbilical  426  ( FIG.  4 A ) from the sled  400  to the electrical interface  125  on the vehicle  100  ( FIG.  1 E ), as shown in  FIG.  7 B . 
     In block  608 , the support arms  416   a - c  are rotated upwardly from the sled  400  to an intermediate position to engage the corresponding support interfaces  127   a - c  on the underside of the vehicle  100  ( FIG.  1 E ), as shown in  FIG.  7 C . The support arms  416   a - c  are releasably coupled to the corresponding interfaces  127   a - c  on the vehicle  100  by suitable “hold and release mechanisms.” Engaging the support arms  416   a - c  with the vehicle  100  provides the stability needed to connect the propellant umbilicals  460   a, b  from the sled  400  to the vehicle  100 , and in block  610 , the propellant umbilicals  460   a, b  are connected from the sled  400  to the corresponding propellant interfaces  124   a, b  on the underside of the vehicle  100  ( FIG.  1 E ), as also shown in  FIG.  7 C . (It should be noted that the propellant umbilicals  460   a, b  are hidden behind the first aft support arm  416   b  in  FIG.  7 C .) 
     Once the propellant umbilicals  460   a, b  have been connected to the vehicle  100 , their connections are verified, and in block  612 , the support arms  416   a - c  are further rotated upwardly enough to raise the vehicle  100  off of its landing gear  126   a - c  (see arrows in  FIG.  7 D ). In block  614 , the landing gear  126   a - c  retract into their respective gear bays on the vehicle  100 , as shown in  FIG.  7 D , and the gear doors  129   a,c  are closed and sealed. In block  616 , all of the umbilical connections between the launch sled  400  and the vehicle  100  (e.g., the connections between the electrical umbilical  426  and the propellant umbilicals  460   a, b  and the vehicle  100 ) are verified to ensure that, for example, the electrical connections have sufficient integrity, and that the propellant connections are properly sealed. Once this has been done, in block  618 , the support arms  416   a - c  are rotated further aft to raise the vehicle  100  to its release angle of attack (AOA) as shown in  FIG.  7 E . In other embodiments, the landing gear  126   a - c  can remain down until the vehicle is raised to its release AOA as shown in  FIG.  7 E , and can then be retracted into their respective gear bays. In one aspect of these embodiments, it should be noted that because the electrical umbilical  426  and propellant umbilicals  460   a, b  are positioned adjacent to the respective support arms  416   a - c  and are approximately the same length as the respective support arms  416   a - c , the length of the umbilicals does not change substantially when the support arms  416   a - c  are rotated from the position shown in  FIG.  7 D  to the position shown in  FIG.  7 E . This avoids the need to provide for a substantial change in length of the umbilicals  426  and  460   a, b  during rotation of the vehicle  100 . 
       FIG.  8 A  is a partially schematic side view illustrating the connections between the vehicle  100  and the forward support arm  416   a  and the electrical umbilical  426 , and  FIG.  8 B  is a partially schematic rear view illustrating the connections between the vehicle  100  and the rear support arms  416   b, c  and the propellant umbilicals  460   a, b , in accordance with embodiments of the present technology. Referring first to  FIG.  8 A , as noted above, the distal end portion of each of the support arms  416   a - c  can be releasably attached to the corresponding support arm interface  127   a - c  by a hold and release mechanism  810  until commanded to release. Although the hold and release mechanisms  810  securely attach the vehicle  100  to the support arms  416   a - c , as described in greater detail below with reference to  FIGS.  9 A- 12 D , the hold and release mechanisms  810  enable the support arms  416   a - c  to rotate relative to the vehicle  100  to move the vehicle  100  relative to the sled  400  as described above with reference to  FIGS.  7 B- 7 D . The hold and release mechanisms  810  can include, for example, one or more mechanical hold-down arms connected to linkages that are pneumatically, hydraulically, and/or electrically controlled to hold an engagement feature (e.g., an engagement surface) on the corresponding interface  127   a - c  until commanded to release in response to a control command. In some embodiments, the hold and release mechanism  810  can be carried on the distal end portion of each support arm  416   a - c  to reduce vehicle weight. In other embodiments, all or a portion of the hold and release mechanism  810  can be incorporated into the vehicle interface  127   a - c . Hold and release mechanisms are known in the art, and in other embodiments, other suitable hold and release systems can be used to attach the support arms  416   a - c  to the corresponding interfaces  127   a - c  until commanded to release. As also shown in  FIG.  8 A , each of the support arm interfaces  127   a - c  can include a corresponding door  802  that automatically closes once the support arm  416   a - c  has been released, to thereby close off the interface and protect it from detrimental aerothermal heating, etc. 
     In some embodiments, the electrical umbilical  426  can be attached to the corresponding electrical interface  125  on the vehicle  100  using any number of suitable connector/receptacle mechanisms known in the art to maintain the electrical connections between the sled  400  and the vehicle  100  until the electrical umbilical  426  is retracted (via, e.g., a lanyard connected to the sled  400 ) for vehicle separation and liftoff. As schematically illustrated in  FIG.  8 A , in some embodiments electrical power and commands are transmitted from the vehicle  100  to the sled  400  via the electrical umbilical  426 , and data/feedback can be transmitted from the sled  400  to the vehicle  100  via the electrical umbilical  426 . In other embodiments, power, commands, data, and/or other electrical information and signals can be transmitted in different directions via the electrical umbilical  426 . As with the support arms  416   a - c , the electrical interface  125  and both of the propellant interfaces  124   a, b  can also include a door  804  that automatically closes upon retraction of the corresponding umbilical. 
     Referring next to  FIG.  8 B , in the illustrated embodiment oxidizer (e.g., subcooled LOX) flows from the sled  400  to the vehicle  100  via the first propellant umbilical  460   a . In one aspect of this embodiment, the first propellant umbilical  460   a  can be further configured to enable oxidizer that is vented from the vehicle oxidizer tank  242  to flow back to the sled  400 . For example, in some embodiments the first propellant umbilical  460   a  can include two separate conduits: one that flows oxidizer (e.g., subcooled LOX) to the vehicle tank  242 , and one that flows vented oxidizer from the vehicle tank  242  back to the sled  400  for re-cooling and re-densifying. As described in greater detail below, by recirculating vented oxidizer back to the launch sled  400  in the foregoing manner, the vehicle oxidizer tank  242  can be maintained at a full level. Similarly, the second propellant umbilical  460   b  can also include two separate conduits: one configured to flow fuel (e.g., Jet-A) from the launch sled  400  to the vehicle  100 , and another to recirculate vented fuel from the vehicle tanks  240   a, b  back to the sled so that the vehicle tanks  240   a, b  are maintained at a full fuel level. The outlets  461   a, b  of the propellant umbilicals  460   a, b  can be releasably attached to the corresponding propellant interfaces  124   a, b  using suitable propellant conduit couplings known in the art. Such couplings can include, for example, pneumatically actuated clamps that maintain the sealed connection between the umbilical outlets  461   a, b  and the corresponding interfaces  124   a, b  until commanded to release. In the illustrated embodiment, the second propellant line  460   b  is also configured to provide one or more electrical umbilical connections between the launch sled and the vehicle  100 . Although it is schematically illustrated that the aforementioned electrical umbilicals extend through the second propellant umbilical  460   b , it will be understood that in other embodiments the electrical connections provided at the second propellant umbilical  460   b  can be positioned outside of the propellant line  460   b  or otherwise in close proximity to the umbilical  460   b.    
       FIGS.  9 A and  9 B  are side and rear views, respectively, of a support arm hold and release mechanism  810  in a first stage of operation, and  FIG.  9 C  is a side view of the hold and release mechanism  810  in a second stage of operation, in accordance with embodiments of the present technology. Referring first to  FIGS.  9 A and  9 B , each of the support arms  416   a - c  on the sled  400  carries a hold and release mechanism  810  on a distal end portion thereof. In some embodiments, each of the hold and release mechanisms  810  includes a fitting  980  (e.g., a “backstop” fitting) that is pivotally attached to the distal end portion of the corresponding support arm  416   a - c  by means of a pivot shaft  982 . Additionally, each of the hold and release mechanisms  810  includes a hold-down arm  984  that is pivotally coupled to the backstop fitting  980  by means of a pivot shaft  985 . The distal end portion of the hold-down arm  984  includes a recessed surface  988  that in operation clamps an attachment fitting  970  of the corresponding vehicle interface  127   a - c  to the backstop fitting  980 . 
       FIGS.  10 A and  10 B  are front and side cross-sectional views, respectively, of the attachment fitting  970  configured in accordance with embodiments of the present technology. Referring to  FIGS.  10 A and  10 B  together, in some embodiments the attachment fitting  970  can be a “U”-shaped fitting having a curved (e.g., rounded convex) clamping surface  1072 . As shown in  FIG.  9 A , the recessed surface  988  on the distal end portion of the hold-down arm  984  has a complementary shape (e.g., rounded concave) to clamp against the surface  1072  of the attachment fitting  970 . 
       FIGS.  11 A and  11 B  are front and side cross-sectional views, respectively, of the backstop fitting  980  configured in accordance with embodiments of the present technology. As these views illustrate, in some embodiments the backstop fitting  980  can include a pocket  989  (e.g., a rectangular-shaped pocket) having side walls  990   a, b  and a back wall  990   c . The pocket  989  is shaped and sized to receive and retain the attachment fitting  970  against downward and aft movement, yet enable the attachment fitting  970  to move forward and upwardly out of the backstop fitting  980  when released by the hold-down arm  984 , as described in greater detail below. 
     As described above, the support arms  416   a - c  are configured to rotate as shown in  FIGS.  7 D and  7 E  after attachment to the corresponding vehicle interfaces  127   a - c . To accommodate this rotation, the hold and release mechanism  810  includes extensible actuators  992   a, b , each having a distal end portion attached to the backstop fitting  980 . The actuators  992   a, b  are configured to extend as the corresponding control arms  416   a - c  rotate aft, thereby causing the corresponding backstop fitting  980  to rotate about the pivot shaft  982  and accommodate rotation of the support arm  416   a - c  relative to the corresponding vehicle attachment fitting  970 . In some embodiments, the actuators  992   a, b  can be hydraulic actuators, pneumatic actuators, electro-mechanical actuators, or other suitable types of actuators, including other nonlinear actuators known in the art. 
     As shown in  FIG.  9 C , when the hold and release mechanism  810  is commanded to release, the hold-down arm  984  rotates away from the attachment fitting  970 , thereby enabling the fitting  970  to move upwardly and forwardly out of the backstop fitting  980  as the vehicle  100  ( FIG.  1 A ) separates away from the sled  400  ( FIG.  4 A ). Movement of the hold-down arm  984  can be controlled by various suitable drive systems known in the art. For example, in some embodiments the hold-down arm  984  can be held in the “hold” position ( FIG.  9 A ) and then moved to the “release” position ( FIG.  9 C ) by a linkage that is hydraulically or pneumatically actuated. In other embodiments, movement of the hold-down arm  984  can be controlled by an electro-mechanical system. 
     The hold and release mechanism  810  described above with reference to  FIGS.  9 A- 11 B  is one example of a suitable hold and release system that can be used with embodiments of the present technology. In other embodiments, other suitable hold and release mechanisms can be used without departing from the present disclosure. For example,  FIGS.  12 A and  12 B  are side and rear views, respectively, of a hold and release mechanism  1210  in a first stage of operation, and  FIGS.  12 C and  12 D  are corresponding side and rear views, respectively, of the hold and release mechanism  1210  in a second stage of operation, in accordance with embodiments of the present technology. Referring first to  FIGS.  12 A and  12 B , the hold and release mechanism  1210  can be referred to as a “clamshell” type hold and release mechanism that utilizes two hold-down members  1274   a, b  to securely attach an end fitting  1270  at each of the vehicle interfaces  127   a - c  to the corresponding support arm  416   a - c . In the illustrated embodiment, the hold-down members  1274   a, b  are arcuate ring segments (or “clamshells”) having an inner diameter configured to encircle a cylindrical crossbar  1272  that extends horizontally on the end fitting  1270 . The proximal end portions of the first and second hold-down members  1274   a, b  are pivotally coupled to a base  1290  by means of corresponding pivot pins  1276   a, b . The distal end portions of the first and second hold-down members  1274   a, b  are releasably secured to the base  1290  in the closed position by a lockpin  1278  that extends through a bore in the base  1290  and through corresponding holes (not shown) in the distal end portions. 
     Turning next to  FIGS.  12 C and  12 D , when the hold and release mechanism  1210  is commanded to release, the lockpin  1278  is retracted from the distal end portions of the first and second hold-down members  1274   a, b  and the hold-down members  1274   a, b  are rotated aft about the corresponding pivot pins  1276   a, b . Movement of the hold-down members  1274   a, b  in the foregoing manner releases the crossbar  1272  and enables the attachment fitting  1270  to move forwardly and upwardly away from the hold and release mechanism  1210  as the vehicle  100  separates from the sled  400 . The hold and release mechanisms described above are provided by way of example only. Accordingly, those of ordinary skill in the art will appreciate that other suitable hold and release mechanisms can be used, consistent with the present disclosure. 
       FIG.  13    is a schematic diagram illustrating a ground based propellant management system  1300  connected to the vehicle  100  and the sled  400  in accordance with embodiments of the present technology. In some embodiments, the propellant management system  1300  includes an oxidizer management system  1302  (e.g., a LOX management system) and a fuel management system  1304  which are operably connected to a controller  1340 . Those of ordinary skill in the art will understand that the oxidizer management system  1302  and the fuel management system  1304  include electronically controlled propellant pumps, valves, actuators, and associated propellant conduits, etc. configured to direct the propellant flows in the manner described herein. The propellant management system controller  1340  includes one or more processors that control operations and functions of the oxidizer management system  1302 , the fuel management system  1304 , and other components of the propellant management system  1300  in accordance with computer-readable instructions stored on a suitable memory. It will be understood that the controller  1340  can include any logic processing unit, such as one or more CPUs, DSPs, APUs, etc. The propellant management system  1300  further includes a heat exchanger  1306  and a liquid nitrogen (LN2) supply  1308 . The LN2 supply  1308  is configured to circulate LN2 through the heat exchanger  1306  to increase the density of LOX flowing through the heat exchanger  1306  by lowering its temperature from the normal boiling point (NBP) of LOX (i.e., −182.96 degrees C.) to the NBP of LN2 (i.e., −195.79 degrees C.). It will be appreciated that although the system of  FIG.  13    is described in the context of a LOX management system, such system could be used with other types of oxidizer. 
     In the illustrated embodiment, the propellant management system  1300  includes a LOX feed line  1310  that can be releasably connected in fluid communication to the sled LOX tanks  452 , and a fuel feed line  1312  that can be releasably connected in fluid communication to the sled fuel tanks  450 . Although the sled  400  includes three LOX tanks  452  and three fuel tanks  450  in some embodiments, the LOX tanks  452  and the fuel tanks  450  are depicted as a single LOX tank and a single fuel tank, respectively, in  FIG.  13    for ease of illustration. The propellant management system  1300  further includes a LOX return line  1318  and a fuel return line  1330 . The LOX return line  1318  can be releasably connected in fluid communication to the sled LOX tanks  452 , the auxiliary LOX tank  458 , and the vehicle LOX tank  242  to enable vented LOX to recirculate through the heat exchanger  1306 . The fuel return line  1330  can be releasably connected to the sled  400  in fluid communication with the vehicle fuel tank  240 . 
       FIG.  14    is a flow diagram of a routine  1400  for loading propellants onto the vehicle  100  and the sled  400  in accordance with embodiments of the present technology. All or portions of the routine  1400  can be performed by ground control computers, the propellant management system controller  1340 , and/or computers onboard the vehicle  100  in accordance with computer readable instructions stored on memory. Referring to  FIG.  14    with reference to  FIG.  13   , in block  1402 , once the vehicle  100  is at its liftoff angle of attack, NBP LOX is circulated through the LN2 heat exchanger  1306 . This increases the density of the LOX by lowering its temperature from LOX NBP (i.e., −182.96 degrees C.), or from about LOX NBP, to LN2 NBP (i.e., −195.79 degrees C.), or to about LN2 NBP. In block  1404 , the LOX management system  1302  transfers the densified LOX from the heat exchanger  1306  to the sled LOX tanks  452  and the auxiliary LOX tank  458  via the LOX feed line  1310 . It will be understood that, in some embodiments the propellant management system  1300  and/or the sled  400  can include one or more propellant pumps and/or associated valves operably connected to the LOX feed line  1310  to facilitate and control the transfer of LOX from the heat exchanger  1306  to the sled LOX tanks  452  and the auxiliary LOX tank  458 . As the LOX in the tanks  452  and  458  warms to its NBP, it expands and is vented from the tanks and returns to the heat exchanger  1306  via the recirculation line  1318  so that the LOX can be re-densified. The heat exchanger  1306  maintains the LOX at the target temperature of, e.g., −195.79 degrees C., or about −196 degrees C., and a target pressure of, e.g., about 17 psia. In other embodiments, the heat exchanger  1306  can maintain the LOX at other target temperatures and/or target pressures. For example, in some embodiments the heat exchanger  1306  (and/or other components of the propellant management system  1300 ) can maintain the LOX at a target pressure less than 17 psia, or greater than 17 psia, such as about 20 psia, or about 40 psia. 
     In block  1406 , once the temperature and pressure of the LOX in the sled tanks  452  and  458  have stabilized at the target temperature and pressure, densified LOX is loaded into the vehicle LOX tank  242  from the sled auxiliary LOX tank  458  via the first propellant umbilical  460   a . It will be understood that, in some embodiments the sled  400  and/or the vehicle  100  can include one or more propellant pumps and/or associated valves operably connected in fluid communication with the first propellant umbilical  460   a  to facilitate and control the transfer of LOX from the sled auxiliary LOX tank  458  into the vehicle LOX tank  242 . As the LOX in the vehicle LOX tank  242  warms and expands, it is vented from the LOX tank  242  through a recirculation umbilical  1320  back to the sled  400 , and then back to the heat exchanger  1306  via the recirculation line  1318 . Although the recirculation umbilical  1320  is depicted as being separate from the first propellant umbilical  460   a  for ease of illustration, it will be understood from the description of the first propellant umbilical  460   a  provided above with reference to  FIG.  8 B  that, in some embodiments, the LOX recirculation umbilical  1320  can be conjoined or otherwise combined with the first propellant umbilical  460   a.    
     In block  1408 , the LOX flowing back to the heat exchanger  1306  is cooled and densified as described above and can be recirculated back to the sled LOX tanks  452 , the sled auxiliary LOX tank  458 , and the vehicle LOX tank  242 , as needed to maintain the tanks in full, or at least nearly full, conditions. In block  1410 , once the LOX in the sled and vehicle tanks has stabilized, the fuel management system  1304  begins circulating fuel through a fuel heat exchanger (not shown) until it reaches its target temperature and pressure (e.g., about 60 degrees F. and about 40 psia). In other embodiments, the fuel heat exchanger can maintain the fuel at other target temperatures and/or target pressures. For example, in some embodiments the fuel heat exchanger (and/or other components of the propellant management system  1300 ) can maintain the fuel at a target pressure less than 40 psia, such as about 17 psia, or about 20 psia, or greater than 40 psia. In block  1412 , the fuel is then loaded into the sled fuel tanks  450  and the sled auxiliary fuel tank  456  via the fuel feed line  1312  and circulated to maintain the fuel at its target temperature and pressure. It will be understood that, in some embodiments the propellant management system  1300  and/or the sled  400  can include one or more propellant pumps and/or associated valves operably connected to the fuel feed line  1312  to facilitate and control the transfer of fuel from the fuel management system  1304  to the sled fuel tanks  450  and the sled auxiliary fuel tank  456 . In block  1414 , fuel then flows from the auxiliary fuel tank  456  on the sled  400  to the vehicle fuel tank  240  via the second propellant umbilical  460   b . In some embodiments, the sled  400  and/or the vehicle  100  can include one or more propellant pumps and/or associated valves operably connected in fluid communication with the second propellant umbilical  460   b  to facilitate and control the transfer of fuel from the sled auxiliary fuel tank  456  into the vehicle fuel tank  240 . The fuel circulates from the vehicle fuel tank  240  back to the sled  400  via a recirculation line  1326 . Although the recirculation line  1326  is illustrated as being separate from the second propellant umbilical  460   b  for ease of illustration, as described above with reference to  FIG.  8 B , in some embodiments the fuel recirculation path from the vehicle  100  to the sled  400  can be through a conduit that is carried by the second propellant umbilical  460   b . The recirculated fuel then flows from the sled  400  back to the fuel management system  1304  via the fuel recirculation line  1330  to maintain the fuel at the target temperature and pressure. Circulation of the fuel through the tanks  450 ,  456  and  240  and the fuel management system heat exchanger as described above ensures that the fuel maintains its target temperature and pressure throughout the system. 
       FIG.  15    is a schematic diagram of a suitable control system architecture for performing all or a portion of the routines described above when the vehicle  100  is connected to the propellant management system  1300 , in accordance with embodiments of the present technology. In the illustrated embodiment, ground control  1500  (e.g., one or more controllers and/or other processing devices executing computer readable instructions and/or responding to user inputs) controls the propellant management system  1300 , ensuring that the temperature, pressure and flow rates of the propellants remain constant, or at least approximately constant, as the propellants flow through the sled  400  and the vehicle  100 . The propellant management system  1300  provides the status of the propellants and other relevant data about the propellant management system (e.g., data feedback) to ground control  1500  and the vehicle  100 . Ground control  1500  also controls operations of the vehicle  100  at this time and, through the vehicle  100 , is able to control the sled  400 . In other embodiments, ground control  1500  can control operation of the sled  400  directly. The sled provides data feedback to the vehicle  100 , which in turn provides data feedback from both the sled and the vehicle  100  to ground control  1500 . Ground control and the vehicle  100  remain in constant communications (e.g., verbal communications from the crew of the vehicle  100  to ground control personnel) throughout the entire process with ground control  1500  providing updates to the crew of the vehicle  100  and the crew providing information to ground control, when necessary. In other embodiments, the communication, control, and/or feedback paths between ground control  1500 , the vehicle  100  and/or the sled  400  can differ from those described above. As further illustrated in  FIG.  15   , and as described in detail above, the propellant management system  1300  provides propellants to both the sled and the vehicle and maintains the propellants at target temperatures and pressures via recirculation. It will be understood by those of ordinary skill in the art that the communications, controls, and data feedback provided between the ground control  1500 , the propellant management system  1300 , the sled  400 , and the vehicle  100  will generally be implemented via wired and/or wireless connections providing digital communications of information, data, control signals, etc. between the respective system controllers. 
       FIG.  16 A  is a flow diagram of a routine  1600   a  for operating the vehicle  100  and the launch sled  400  during a takeoff run;  FIG.  16 B  is a flow diagram of a routine  1600   b  for confirming safe liftoff conditions prior to release of the vehicle  100  from the sled  400 ;  FIG.  16 C  is a flow diagram of a routine  1600   c  for liftoff of the vehicle  100  from the sled  400 ; and  FIG.  16 D  is a flow diagram of a routine  1600   d  for aborting liftoff of the vehicle  100  from the sled  400 , in accordance with embodiments of the present technology.  FIGS.  17 A- 17 D  are a series of schematic diagrams illustrating operation of the vehicle  100  and the launch sled  400  at various stages of the launch sequences described in  FIGS.  16 A- 16 D , in accordance with embodiments of the present technology. Referring first to the routine  1600   a  of  FIG.  16 A , in block  1602 , once all of the pre-flight checkouts for the vehicle  100  have been completed and the vehicle is ready for takeoff, ground control stops propellant flow from the propellant management system  1300  ( FIG.  13   ) to the sled  400  and then disconnects the propellant management system  1300  from the sled  400 , as shown in  FIG.  17 A . In block  1604 , the sled propellant system continues to maintain the sled and vehicle propellant conditions until engine ignition by recirculation of LOX and fuel from the vehicle LOX tank  242  and the vehicle fuel tank  240 , respectively, to the sled LOX tanks  452  and  458 , and the sled fuel tanks  450  and  456 , respectively, as also shown in  FIG.  17 A . 
     In block  1606 , ground control commands the sled and vehicle engines  404   a - c  and  120   a - c , respectively, to ignite and throttle up to 100 percent, as shown in  FIG.  17 B . In block  1608 , all the engines  120   a - c  on the vehicle  100  are then set to “tripped flow” for optimal, or at least near optimal, low altitude performance. In some embodiments, setting the engines  120   a - c  to tripped flow changes the area ratio of the engines from, e.g., 60:1 to, e.g.,  33 : 1 . Together, the vehicle engines  120   a - c  and the sled engines  404   a - c  provide thrust to propel the sled  400  and the vehicle  100  down the launch rails  410   a - c  ( FIG.  4 A ) for takeoff. In block  1610 , throughout the entire sled run, the sled LOX and fuel tanks  452  and  450 , respectively, feed propellants to the sled engines  404   a - c , while the auxiliary LOX and fuel tanks  458  and  456 , respectively, refuel the vehicle LOX and fuel tanks  242  and  240 , respectively. As described above and elsewhere herein, refueling of the vehicle propellant tanks in the foregoing manner enables the vehicle  100  to lift off from the sled  400  with full, or at least approximately full, propellant tanks, thereby eliminating, or at least greatly reducing, the liftoff dry weight penalty associated with conventional space launch vehicles. 
     In some embodiments, the vehicle  100  is in the takeoff angle of attack position shown in  FIG.  7 E  during the takeoff run. In other embodiments, instead of being positioned at the takeoff angle of attack during the entire takeoff run, the vehicle  100  can begin the takeoff run at a lower angle of attack (such as the position shown in  FIG.  7 D ) to reduce drag and increase acceleration, and then the support arms  416   a - c  can rotate aft to increase the angle of attack of the vehicle  100  just prior to liftoff and separation from the sled  400 . 
     If any anomalies are detected during any part of the sled run after block  1610 , the routine proceeds to block  1616  and initiates a refused takeoff sequence. Alternatively, if no anomalies are detected, the routine proceeds to block  1612  and, at or just before reaching takeoff speed, the propellant flow from the sled  400  to the vehicle  100  is stopped, the first and second propellant umbilicals  460   a  and  460   b  are disconnected and retracted from the vehicle  100 , and the propellant doors  804  ( FIG.  8 A ) on the vehicle  100  are closed, as depicted by  FIG.  17 C . From this point forward, the vehicle engines  120   a - c  are using propellant solely from the vehicle LOX and fuel tanks  252  and  240 , respectively. If anomalies are detected during or after the separation process of block  1612 , the routine proceeds to block  1616  and initiates the refused takeoff sequence. If not, then the routine proceeds to block  1614  and the vehicle and sled systems (e.g., the respective controllers) check their respective operating systems to confirm that safe liftoff conditions have been met. 
     Referring next to  FIG.  16 B , the routine  1600   b  starts at block  1614  from the routine  1600   a  above. In block  1614 , just prior to vehicle separation from the sled  400 , the vehicle  100  and the sled  400  perform a number of system checks to confirm that safe liftoff conditions are met. The system checks can be run concurrently, and can include, for example, confirming that all engines on the vehicle  100  and the sled  400  are operating nominally (block  1620 ), all flight critical systems are operating nominally (block  1622 ), and the lift on the vehicle is equal to or greater than 1.1 times the gross takeoff weight of the vehicle (block  1624 ). If one or more of the system checks fail or are unsuccessful, the routine proceeds to block  1626  and executes the refused takeoff sequence. Conversely, if all of the system checks are successful, the routine proceeds to block  1628  and the hold and release mechanisms on the sled support arms  416   a - c  are commanded to release the vehicle  100  as depicted in  FIG.  17 D . 
     Referring next to  FIG.  16 C , the routine  1600   c  starts at block  1628  from the routine  1600   b  above. In block  1628 , once the hold and release mechanisms on the sled support arms have been commanded to release the vehicle  100 , the routine proceeds to block  1630  and the vehicle  100  separates and lifts off from the sled  400 . In block  1632 , the electrical/data umbilicals (e.g., the electrical umbilical  426  ( FIG.  4 A )) are detached from the vehicle interface  125  ( FIG.  1 E ) and retracted (e.g., retracted onto the forward support arm  416   a  ( FIG.  4 A )), and in block  1634  all remaining doors (e.g., the support arm doors  802  and the electrical/propellant umbilical doors  804  ( FIG.  8 A )) are closed. At this time, the vehicle controller  140  can confirm that all of the electrical and propellant umbilicals have been disconnected from the vehicle  100  and all of the associated doors on the vehicle  100  have been closed. In block  1636 , the sled engines  404   a - c  are shut down and in block  1638  the sled slows down and comes to a stop when it encounters the water barriers at the end of the sled run. 
     Turning next to  FIG.  16 D , as noted above  FIG.  16 D  is a flow diagram of a routine  1600   d  for executing a refused takeoff sequence in accordance with embodiments of the present technology. In block  1640 , the routine begins when anomalies are detected at any point during the sled run. In block  1642 , all engines on the sled  400  and the vehicle  100  are immediately commanded to shut down, and in block  1644  all sled and vehicle systems are commanded to safe conditions. Such safe conditions can include, for example, opening tank vents on the LOX tanks on both the sled  400  and the vehicle  100 . In block  1646 , the sled  400  with the vehicle  100  still mounted thereon slows and safely brakes to a stop when the sled  400  encounters the water barriers at the end of the sled run. 
       FIG.  18 A  is a schematic diagram of a suitable control system architecture  1800   a  for performing all or a portion of the routines  1600   a ,  1600   b , and  1600   d  described above after the vehicle  100  and the sled  400  have detached from the propellant management system  1300 , in accordance with embodiments of the present technology, and  FIG.  18 B  is a schematic diagram of a suitable control system architecture  1800   b  for performing all or a portion of the routine  1600   c  after the vehicle  100  has detached from the sled  400 , in accordance with embodiments of the present technology. Referring first to  FIG.  18 A , when the vehicle  100  and the sled  400  have disconnected from the propellant management system  1300 , the vehicle  100  is in control of the sled  400  and provides both electrical power and control actuation commands to the sled  400 . The sled  400  provides data feedback on the commanded controls from the vehicle  100  and on the status of the propellant as it continues flowing from the sled  400  into the vehicle  100 . The vehicle  100  provides data feedback on all systems in both the sled  400  and the vehicle  100  to ground control  1500 . In some embodiments, ground control  1500  is no longer able to control the vehicle  100  at this point, but both maintain constant communications (e.g., wireless communications) throughout the entire process, providing information to the other when necessary. In other embodiments, ground control  1500  can provide all, or a portion, of the control inputs for the vehicle  100  after separation from the sled  400 . Accordingly, it will be understood that ground control  1500  includes suitable communications systems for wireless transmission of communications, command signals, and telemetry to and from the vehicle  100 . 
     Turning next to  FIG.  18 B , once the vehicle  100  has disconnected from the sled  400 , the vehicle  100  is able to provide data feedback from its onboard systems to ground control  1500 , but ground control  1500  is not able to directly control the vehicle  100 . Instead, vehicle guidance, navigation, control, systems management, etc. is performed by the vehicle controller  140  in accordance with flight control software and trajectory optimization code tasks, as described in greater detail below. As noted above, however, in other embodiments ground control  1500  can provide all or a portion of the control commands and/or input necessary for vehicle guidance, navigation, control, and/or systems management. Throughout the flight, ground control  1500  and the vehicle  100  can maintain constant communication, providing information to the other when necessary. 
       FIG.  19    is a partially schematic diagram illustrating various phases in a flight sequence of the vehicle  100  in accordance with some embodiments of the present technology. In a takeoff phase  1901 , the sled  400  is mounted to the launch rails  410   a - c  as described above with reference to  FIG.  4 A , and the vehicle  100  is mounted to the sled  400  as described above with reference to  FIG.  7 E . On a typical flight, the vehicle  100  may carry a payload of about 5,000-7,500 lbs. and a crew of five. Prior to takeoff, the vehicle engines  120   a - c  ( FIG.  1 A ) and the sled engines for  404   a - c  are ignited and brought up to full thrust using propellant from the sled  400  as described above. If needed, the sled  400  can be held in place on the rails  410   a - c  as the engines are brought up to full thrust using a sled braking system or a hold and release mechanism. During the takeoff run, the additional thrust from the sled  400  boosts the vehicle acceleration, and the use of sled propellants instead of vehicle propellants enables the vehicle  100  to take off fully loaded with propellant. At or near takeoff speed, the electrical and propellant umbilicals disconnect from the vehicle  100  and the vehicle  100  releases from the support arms  416   a - c  and enters a pull up phase  1902 . 
     In some embodiments, the vehicle  100  can achieve a takeoff speed of from about 400 mph to about 500 mph, or about 436 mph (0.7 Mach), in about 20 seconds after traveling down the rails  410   a - c  a distance of from about 4,500 ft. to about 6,000 ft., or about 5,182 ft. In some embodiments, the vehicle  100  and crew may experience relatively low dynamic forces during takeoff, with maximum accelerations ranging from about 1 g to about 2 g&#39;s, or about 1.42 g&#39;s. The foregoing launch parameters are illustrative of some embodiments of the present technology. In other embodiments, the vehicle  100  can achieve different takeoff speeds in different takeoff run distances, and resulting in different maximum acceleration levels, without departing from the present disclosure. 
     As described above with reference to  FIG.  4 A , the sled  400  can include a braking system  455 . In some embodiments, the braking system  455  can enable the sled  400  to decelerate, with the vehicle  100  attached to the sled  400 , from the takeoff speed to zero in a distance of from about 3,000 ft. to about 3,500 ft., or about 3,200 ft., resulting in a maximum deceleration of from about 1 g to about 3 g&#39;s, or about 2 g&#39;s. This sled braking feature can enable the sled  400  to perform a refused takeoff at any point prior to takeoff if necessary for mission safety. 
     After the vehicle  100  lifts off from the sled  400 , the vehicle flight path is controlled by operation of the aerodynamic control surfaces described above with reference to  FIG.  1 A  and/or gimballing the engines  120   a - c . The pull up phase  1902  can be relatively gentle and permit a low velocity turn to a target azimuth that provides orbital inclination flexibility. In some embodiments, the maximum accelerations the vehicle  100  experiences during the pull up phase  1902  can be from about 1.2 g&#39;s to about 2 g&#39;s, or about 1.7 g&#39;s. After the pull up phase  1902 , the vehicle  100  enters an ascent phase  1903  in which the vehicle may be limited to maximum accelerations of, for example, about 3 g&#39;s to maintain crew/passenger comfort. Vehicle directional control during all or a portion of the ascent phase can be provided, or at least supplemented, by engine thrust vector control and/or engine gimbaling. During the ascent phase  1903 , the vehicle can experience a maximum dynamic pressure (max Q) of from about 1,000 psf to about 1,100 psf, or about 1,080 psf. In some embodiments, at the end of the ascent phase  1903  the vehicle  100  will be traveling at a speed of from about 15,000 mph to about 20,000 mph, or about 17,560 mph or more, and the vehicle will be at an altitude of from about 275,000 ft. to about 325,000 ft., or about 300,000 ft. or more. 
     Turning next to  FIG.  20   , this Figure is a flow diagram of a routine  2000  for performing an ascent sequence of the vehicle  100  in accordance with embodiments of the present technology. Upon separation from the sled  400  in block  2002 , the vehicle  100  maintains both the launch azimuth and a preset rate of climb until the aerodynamic effects from separation are dampened. In block  2004 , the vehicle continues the ascent and banks to the target azimuth as commanded by the vehicle controller  140  ( FIG.  1 A ) in accordance with flight control software executing a trajectory optimization code. In block  2006 , when the optimal altitude has been reached for changing the area ratio of the nozzles of the engines  120   a - c  (e.g., typically about 32,000 ft.), the tripped flow is turned off and the engine area ratio is increased from, e.g., about 33:1 to, e.g., about 60:1. In some embodiments, an area ratio of 60:1, or at least approximately 60:1, can provide optimal, or near optimal, high altitude flight performance of the vehicle engines  120   a - c.    
     In block  2008 , all three of the engines  120   a - c  maintain 100 percent throttle until, in block  2010 , a preset axial acceleration limit is reached. In some embodiments, the preset axial acceleration limit can be 3 g&#39;s. In other embodiments, other preset axial acceleration limits can be used that are higher or lower than 3 g&#39;s. Once the axial acceleration limit has been reached, all the engines  120   a - c  begin throttling down to maintain the axial acceleration at or below the limit. When the lower throttle limit is reached on the engines  120   a - c  (typically about 50 percent throttle), the outboard engines  120   a  and  120   c  are shut down, and the center engine  120   b  ( FIG.  1 A ) is throttled up to 100 percent. In block  2014 , when the preset axial acceleration limit is again reached, the center engine  120   b  begins throttling down until it reaches its lower throttle limit (e.g., about 50 percent). In block  2020 , upon vehicle insertion into the target orbit, the vehicle center engine  120   b  shuts down. Final adjustments to the orbit of the vehicle  100  can then be performed using the OMS engines  122   a, b  ( FIG.  1 A ). 
     As noted above, all or a portion of the vehicle trajectory and control during the ascent sequence of  FIG.  20   , and/or other portions of the flight of the vehicle  100 , can be controlled by the vehicle controller  140  in response to execution by the processer  142  of computer-readable instructions stored on non-volatile memory  144  ( FIG.  1 A ). The computer-readable instructions executed by the processor  142  can include flight control software and trajectory optimization code. In some embodiments, the trajectory optimization code can include tasks for flight profile optimization and engine management. For example, in some embodiments the trajectory optimization code tasks can include optimizing the vehicle angle of attack profile and the vehicle bank angle profile to achieve the desired ascent trajectory. Additionally, in some embodiments the code tasks can also include optimizing engine control to determine, for example, when to throttle down the engines, shut off the outboard engines, change the engine area ratios, etc. The trajectory optimization code tasks can also include maintaining flight path constraints. For example, in some embodiments this can include maintaining normal acceleration (i.e., acceleration along an axis perpendicular to the longitudinal axis of the vehicle and extending from the bottom of the vehicle to the top of the vehicle) at less than or equal to 2.5 g&#39;s, maintaining axial acceleration at less than or equal to 3 g&#39;s, and maintaining vehicle dynamic pressure at less than or equal to 1,200 psf. The foregoing flight path constraints are examples of some suitable constraints for some embodiments of the present technology. Accordingly, other embodiments can utilize other flight path constraints. In addition to the foregoing, the trajectory optimization code tasks can further include targeting the desired final conditions upon insertion into the target orbit. Such conditions can include, for example, maximizing the final vehicle weight and achieving preset orbital targets. In some embodiments, the preset targets can include a perigee altitude of 50 nautical miles, an apogee altitude of 100 nautical miles, and an orbital inclination of 51.6 degrees. The foregoing orbital parameters are provided by way of example, and in other embodiments the trajectory optimization code can be tasked with achieving other final conditions, orbits, etc. 
     Returning now to  FIG.  19   , in an orbital phase  1904  the vehicle  100  can engage in various orbital operations including, for example, transfer of crew and/or cargo to on-orbit stations. Typical missions on orbit could be from about 3 to 5 days. Once orbital operations are complete, the vehicle  100  can move to a tail-first orientation using the RCS thrusters  136  described above with reference to  FIGS.  1 D and  1 E . Next, the vehicle  100  can ignite the OMS engines  122   a, b  ( FIG.  1 A ) to reduce orbital speed and deorbit, thereby entering a reentry phase  1905  in which the vehicle  100  descends through the Earth&#39;s atmosphere. In some embodiments, the vehicle  100  can have a weight of about 60,000 lbs. and be at an altitude of about 400,000 ft. at a reentry start point  1905   a . In some embodiments, the angle of attack and the bank angle of the vehicle  100  can be optimized during reentry so that the vehicle  100  will be at an altitude of about 50,000 ft. and travelling at a speed of about 0.6 Mach at a reentry end point  1905   b  that is about 3,000 nautical miles from the reentry start point  1905   a . Accelerations during the reentry phase  1905  are relatively low and can range from a maximum acceleration of from about 1.5 g&#39;s to about 2.5 g&#39;s. During a final glide phase  1906  and subsequent landing phase  1907 , the vehicle speed and descent rate can be at least generally similar to the speed and descent of a conventional commercial jet aircraft. For example, in some embodiments the vehicle  100  can land at a speed of from about 120 mph to about 160 mph, or about 140 mph, and at a sink rate of about 10 feet per second. Additionally, the vehicle  100  can land on a standard runway  1580 . Although the foregoing reentry and landing parameters of some embodiments are provided herein by way of example, in other embodiments the vehicle  100  and missions thereof can have other reentry and landing parameters. 
     In the event the vehicle  100  lands at a runway that does not have a suitable launch sled, the vehicle  100  can be moved to a runway that does have a launch sled using a number of different methods. For example, in some embodiments the vehicle can be towed through the air to a new runway by a tow aircraft. In other embodiments, the vehicle  100  can include provisions for jet engines that can be temporarily installed on the wings  104   a, b  ( FIG.  1 A ) to enable the vehicle  100  to fly to the new runway under its own power. 
     If an irrecoverable emergency arises during any phase of flight, the vehicle  100  can execute an abort phase  1908 . In the abort phase  1908 , the crew cabin  112  separates from the rest of the vehicle  100  as described above with reference to  FIG.  3 A . Immediately after separation, the high-power escape thrusters  360   a, b  propel the crew cabin  112  safely away from the rest of the vehicle  100 , and the recovery chute  364  is deployed so that the crew cabin  112  can descend to a safe landing. The emergency parachute landing system is configured to bring the crew cabin  112  down safely on land or in water, and the cabin  112  has provisions that permit crew survival for an extended period of time if rescue is delayed. 
       FIG.  21    is a chart listing example types of mission aborts and engine failures/degradations that the vehicle  100  could experience, in accordance with embodiments of the present technology. In some embodiments, there are four main types of mission aborts and two main types of engine failures. The least impactful of the aborts is a “To Orbit Abort”  2102 . This abort condition arises when the vehicle  100  will achieve orbit safely but will fall short of target performance. To successfully abort the mission under this condition, the vehicle  100  will continue the mission until orbit is achieved, and then mission continuation and landing options will be assessed. A slightly more impactful abort is a “Once Around Abort”  2104 . This abort condition arises when the vehicle  100  will not achieve a stable orbit. To abort this condition, the vehicle  100  will perform one orbit around the Earth and return safely to the launch site. The second most impactful abort is a “Down Range Abort”  2106 . This abort condition arises when the vehicle  100  is unable to perform one orbit around the Earth as required by the Once Around Abort  2104 . This abort sequence calls for the vehicle  100  to perform an emergency landing at one of a preset downrange emergency landing sites (e.g., a landing strip or runway). The most impactful abort is a “Return to Launch Site Abort”  2108 . This abort condition arises when the vehicle performance falls short of the conditions for the Down Range Abort  2106 . In a Return to Launch Site Abort, the vehicle  100  will immediately return to and land at the launch site. In some embodiments, the propellant tanks on the vehicle  100  can include one or more drain valves configured to rapidly discharge the propellant from the tanks prior to landing to ensure that the vehicle  100  lands with empty, or near empty propellant tanks during an abort. Landing with empty, or near empty propellant tanks enables the landing gear  126   a - c  to be substantially lighter than would otherwise be required for landing the vehicle  100  with full, or near full propellant tanks. 
     In some embodiments, the most impactful of the engine failures is an “Outboard Engine Failure”  2110  resulting from the performance of one or both of the outboard engines  120   a, c  ( FIG.  1 A ) degrading to the point of failure. To address this issue, the vehicle  100  executes a sequence that commands both the outboard engines  120   a, c  to shut down and the center engine  120   b  to increase throttle to maximum thrust. The least impactful of these failures is an “Outboard Engine Performance Degradation”  2112 . This failure condition is due to the performance of one of the outboard engines  120   a  or  120   c  degrading, but not to the point of the outboard engine failure condition  2110 . To address this issue, the vehicle  100  executes a sequence that commands the opposing outboard engine to throttle down to match the failed engine&#39;s performance. Throttling the opposing engine in this manner balances the thrust from the outboard engines  120   a, c  and avoids exceeding the yaw control limit of the vehicle  100 . 
       FIG.  22    is a flow diagram of a routine  2200  for executing an outboard engine anomaly abort sequence, in accordance with embodiments of the present technology. In block  2202 , the routine starts when an anomaly in performance of one of the outboard engines  120   a, c  ( FIG.  1 A ) is detected. If the anomaly is an engine failure, then the routine proceeds to block  2204  and performs the outboard engine failure sequence described above with reference to  FIG.  21   . Conversely, if an engine performance degradation is detected in one of the outboard engines, then the routine proceeds to block  2206  and performs the outboard engine performance degradation sequence described above with reference to  FIG.  21   . In either situation, after execution of the appropriate sequence, the routine proceeds to block  2208 . In block  2208 , the vehicle initiates the To Orbit Abort sequence  2102  of  FIG.  21    and, if the To Orbit Abort condition is met (e.g., the vehicle will achieve orbit safely, but will fall short of target performance), then the routine proceeds to block  2210  and the vehicle  100  executes the To Orbit Abort sequence outlined in  FIG.  21    (success). Conversely, if the To Orbit Abort condition is not met (failure), then the routine proceeds to block  2212 . In block  2212 , the vehicle initiates the Once Around About sequence  2104  of  FIG.  21    and, if the Once Around Abort condition is met, then the vehicle  100  executes the appropriate abort sequence (e.g., perform one orbit around Earth and return safely to the launch site) (success). If the Once Around Abort condition is not met (failure), then the routine proceeds to block  2214 . In block  2214 , the vehicle initiates the Down Range Abort sequence  2106  of  FIG.  21    and, if the Down Range Abort condition is met, then the vehicle  100  executes the appropriate abort sequence (success). Conversely, if the Down Range Abort condition is not met (failure), then the routine proceeds to block  2216  and performs the Return to Launch Site Abort sequence  2108  of  FIG.  21    and executes the appropriate abort sequence (e.g., immediately return to and land at the launch site). 
     The flow diagram of  FIG.  22    and the other flow diagrams described herein depict processes used in some embodiments of the present technology. These flow diagrams may not show all functions or exchanges of data, but instead they provide an understanding of commands, data, and/or information exchanged under some embodiments of the systems. Those of ordinary skill in the relevant art will recognize that some functions or exchange of commands and data may be repeated, varied, omitted, or supplemented, and other (less important) aspects not shown in the flow diagrams may be readily implemented. Each of the steps depicted in the flow diagrams described herein can itself include a sequence of operations that need not be described herein. Those of ordinary skill in the art can create source code, microcode, program logic arrays, etc. or otherwise implement the disclosed technology based on the flow diagrams and the Detailed Description provided herein. Such routines are preferably stored in non-volatile memory, e.g., the memory  144  that forms part of the vehicle controller  140  ( FIG.  1 A ). 
       FIG.  23 A  is a block diagram of a suitable computing environment  2300   a  in which the vehicle controller  140  can implement the various sequences and routines described in detail above. In the illustrated embodiment, the vehicle  100  includes a display  2310  and a user interface  2312  which are operably connected to the controller  140 . The display  2310  can include one or more conventional display devices (e.g., LCD displays, LED displays, etc.) for providing textual, graphical, and other information to users (e.g., vehicle crew). The user interface  2312  can include any suitable user interface devices and tools including, for example, touchscreens, keyboards, keypads, joy sticks, graphical user interfaces, etc. In one aspect of the present technology, the vehicle controller  140  can include or access a number of on-board software applications. For example, in the illustrated embodiment the environment  2300   a  includes a guidance, navigation, and control (GNC) application  2302 , a systems management (SM) application  2304 , and a vehicle check out (VCO) application  2306 . The GNC application  2302  is configured to determine flight parameters, such as the position, velocity, and attitude of the vehicle  100  during flight. The GNC application  2302  also receives and manages various outputs from vehicle sensors (e.g., airspeed sensors, altitude sensors, acceleration sensors, pressure sensors, temperature sensors, etc.) and displays the output values to the vehicle crew via the display  2310  and to ground control  1500  via one or more associated displays. In addition, the GNC application  2302  also manages the majority of subsystems aboard the vehicle, such as the avionic subsystems, throughout the entirety of the vehicle flight from after sled separation to vehicle landing. The SM application  2304  manages and controls the remainder of the vehicle subsystems that are not controlled by the GNC application  2302 , such as the payload subsystems, etc. Additionally, the SM application  2304  is also configured to identify anomalies/errors mid-flight and display them to both the vehicle crew and the ground control crew. The VCO application  2306  manages and controls all subsystems (e.g., the avionics subsystems) during their initialization process (which can occur during the vehicle propellant loading sequence described above). The VCO application  2306  also performs all of the ground and in-flight checkouts for the vehicle systems and subsystems including, for example, determining if safe separation conditions are met, if engine performance conditions are met, etc. The VCO application  2306  also processes ground control commands when the vehicle is connected to the propellant management system  1300  ( FIG.  13   ). The vehicle controller  140  can receive inputs from the vehicle crew via the user interface  2312 , from ground control when the vehicle is attached to the propellant management system  1300 , and through data feedback from the GNC application  2302 , the SM application  2304 , and/or the VCO application  2306 . 
     By way of an example implementation of the environment  2300   a , the vehicle controller  140  utilizes the GNC application  2302 , the SM application  2304 , and the VCO application  2306  to perform and execute the mission tasks for the vehicle  100  as described in detail above. The system applications  2302 ,  2304  and  2306  process these tasks ( 2308 ) and, when the processing is successful, a task is executed ( 2314 ) and data feedback is sent back to the vehicle controller  140 . Conversely, when the processing results in a failure, the task is corrected ( 2316 ) and reperformed until successful. The data feedback can be displayed on both the vehicle user interface  2312  or display  2320 , as well as one or more user interfaces associated with ground control  1500 . 
       FIG.  23 B  is a block diagram of a suitable computing environment  2300   b  in which the sled controller  440  can implement one or more of the sled routines described in detail above. In the illustrated embodiment, the sled controller  440  includes and/or accesses a number of software applications, including, e.g., a sled control (SC) application  2320 , a sled systems management (SSM) application  2322 , and a sled checkout (SCO) application  2324 . In operation, the sled controller  440  receives inputs from the vehicle crew via the vehicle controller  140 . The sled controller  440  can also receive inputs from ground control  1500  via the vehicle controller  140  when the vehicle  100  is attached to the propellant management system  1300  ( FIG.  13   ). The sled controller  440  can also receive data feedback from the SC application  2320 , the SSM application  2322 , and the SCO application  2324 . The sled controller  440  uses the three software applications to perform specific tasks in accordance with input and commands received from the vehicle computer  140  and/or ground control  1500 . Such tasks can include, for example, igniting the sled engines  404   a - c , moving the support arms  416   a - c , controlling sled propellant flow, etc. These tasks are processed ( 2326 ) and, if the processing is successful, the tasks are executed ( 2330 ) and feedback data is sent back to the sled controller  440 . Alternatively, if the task processing results in a failure, the task is corrected ( 2328 ) and reperformed until it has been successfully executed. Data feedback to the sled controller  440  can also be sent to the vehicle controller  140  and displayed to both the vehicle crew (via, e.g., the display  2310 ) and ground control  1500  when the sled  400  is attached to the propellant management system  1300 . 
       FIG.  24 A  is a schematic diagram of the vehicle LOX tank  242  ( FIG.  2   ), and  FIG.  24 B  is a graph illustrating various pressures associated with the LOX tank  242  as a function of time, in accordance with embodiments of the present technology. Oxygen is gaseous during normal vehicle operating conditions and liquid at temperatures below its Normal Boiling Point (NPB) of −182.96 degrees C. Since oxygen boils at such a low temperature, it has a high vapor pressure at normal vehicle operating conditions and is typically kept in heavy, round tanks capable of withstanding relatively high pressures (e.g., pressures over 20 psig) in conventional launch vehicles. In one aspect of the present technology, however, the LOX in the vehicle  100  is subcooled to a temperature of, e.g., −195.79 degrees C., or about −196 degrees C., to reduce the vapor pressure to less than 4 psig (e.g., 2-3 psig) across the walls of the LOX tank  242 . As a result of this relatively low pressure, the fuselage/wing LOX tank  242  of the vehicle  100  can be shaped for aerodynamic efficiency and designed to support flight loads, without requiring a rounded, pressure stabilized design as typically found in conventional launch vehicle LOX tanks. This feature also enables the use of lightweight composite materials, which lowers the structural tank weight, and provides a higher LOX load than NPB LOX because the LOX density increases with the lower temperature. The increased LOX load improves vehicle performance by increasing the amount of propellant that can be loaded into a given tank volume, improving the mass fraction of the flight vehicle  100 . 
     Referring to  FIG.  24 A , in the illustrated embodiment the LOX tank  242  can be vented outward through a first relief valve  2472  during vehicle ascent. Additionally, the LOX tank  242  can also include a second relief valve  2474  that enables the LOX tank  242  to vent inward from the atmosphere to prevent negative pressure tank collapse on reentry. Desiccant canisters  2476  can be installed at the vent inlet to prevent ingestion of moisture during ground operations. During ground operations, the LOX will be stored in the LOX tank  242  at the lowest vapor pressure possible (or approaching the lowest vapor pressure possible), resulting in a pressure differential across the tank walls of less than 3 psig. During engine operation, the LOX tank  242  can receive pressurization gas from a heat exchanger  2470  coupled to one or more of the main engines  120   a - c . A boost pump (not shown in  FIG.  24 A ) can be used to increase the LOX pressure to, for example, about 40 psia to meet inlet conditions for the turbo pumps associated with the engines  120   a - c.    
     Before takeoff, the LOX tank  242  can be filled and pressurized using ground-based sources of subcooled LOX (e.g., the propellant management system  1300  of  FIG.  13   ). By way of example, in some embodiments the LOX tank  242  can receive LOX at a temperature of about −196 degrees C. (i.e., about −320 degrees F.) from a LOX densification unit. During engine operation, pressure can be maintained in the LOX tank  242  by adding vaporized propellant gasses supplied by the heat exchanger  2470  on the main engine  120 . In some embodiments, the pressure differential across the tank walls will be maintained in a range from about 2 to 3.5 psig during ascent to prevent boiling of the LOX, and to maintain tank pressure above local ambient (i.e., 14.7 psia or less) to prevent tank buckling due to a negative pressure differential. During ascent and on orbit, residual ullage gasses vent through the first relief valve  2472 . During reentry, the LOX tank  242  vents to the atmosphere via the second relief valve  2474  to prevent negative pressure tank collapse. 
     Operation of the LOX tank  242  as described above is reflected by the graph  2478  shown in  FIG.  24 B . In the graph  2478 , pressure in psi is measured along a vertical axis  2480 , and time in seconds is measured along a horizontal axis  2482 . A first plot line  2484  illustrates the atmospheric pressure during the ascent phase of flight, a second plot line  2486  illustrates tank internal pressure (absolute pressure) during this phase of flight, and a third plot line  2487  illustrates tank differential pressure during this phase of flight. As the first plot line  2484  illustrates, the atmospheric pressure drops from about 14.7 psia at launch to essentially zero during ascent. As shown by the second plot line  2486 , the tank internal pressure follows this curve relatively closely to maintain a positive pressure differential of about 3.5 psig at all times during the launch and ascent phases of flight, as shown by the third plot line  2487 . 
       FIG.  25 A  is a schematic diagram of the left-wing fuel tank  240   a  ( FIG.  2   ), and  FIG.  25   b    is a graph  2596  illustrating various pressures associated with the fuel tank  240   a  during vehicle ascent, in accordance with embodiments of the present technology. Although the foregoing description refers to the left-wing fuel tank  240   a , it will be understood that the description applies equally to the right-wing fuel tank  240   b . Referring to  FIG.  25 A , as noted above in some embodiments the vehicle  100  will use Jet-A as the fuel for the vehicle main engines  120   a - c . Jet fuel is abundant, inexpensive, and has a naturally low vapor pressure across vehicle operating conditions which allows it to be stored inside aerospace vehicle wings, fuselages, etc. of virtually any shape. In the illustrated embodiment, the fuel tank  240   a  includes a first relief valve  2592 , a second relief valve  2594 , and a boost pump  2590 . As with the LOX tank  242  described above with reference to FIG.  24 A, during engine operation the fuel tank  240   a  can maintain positive pressure by the addition of vaporized propellant gasses supplied from a heat exchanger on one or more of the vehicle engines  120  (not shown). In orbit, residual ullage gasses can be vented from the fuel tank  240   a  through the first relief valve  2592 . During reentry, the fuel tank  240   a  can be vented to atmosphere via the second relief valve  2594  to prevent negative pressure collapse. Although not shown, desiccant canisters can be installed at the vent inlet to the relief valve  2594  to prevent ingestion of moisture during ground operations. 
     Referring next to  FIG.  25 B , pressure in psi is measured along a vertical axis  2595 , and time in seconds is measured along a horizontal axis  2597 . A first plot line  2591  illustrates the atmospheric pressure as a function of time during vehicle ascent, and a second plot line  2593  illustrates the internal (absolute) pressure of the fuel tank  240   a  during this phase of flight. The differential tank pressure across the tank wall is illustrated by a third plot line  2598 . In this embodiment, the first relief valve  2592  is open and the fuel tank  240   a  vents so that the internal pressure (second plot line  2593 ) is essentially equivalent to the atmospheric pressure (first plot line  2591 ) for the initial portion of vehicle ascent. At about 47 seconds after takeoff, the first relief valve  2592  closes. At this time, the vehicle will be at an altitude of about 13,000 ft. and the atmospheric pressure will be about 9 psia. As vehicle ascent continues, the tank internal pressure is allowed to build relative to the atmospheric pressure as shown by a comparison of the first plot line  2591  to the second plot line  2593 . As a result of this pressure differential, the fuel tank  240   a  has a positive pressure differential across the tank walls of from about 3 psig to about 1.4 psig at the end of the vehicle ascent phase, as shown by the third plot line  2598 . 
       FIGS.  26 A and  26 B  are top and bottom isometric views, respectively of the vehicle  100 . In some embodiments, the vehicle primary structure, external surfaces, etc. (including, for example, the wings  104   a, b , the fuselage  102 , the vertical stabilizers  110   a, b , etc.) can be constructed from lightweight, durable composite materials. The composite materials can include graphite and an epoxy matrix including, for example, a polyurethane that is resistant to micro cracking and oxygen infusion. In some embodiments, portions of vehicle  100  can also be constructed from metal including, for example, aluminum, titanium, stainless steel, etc. Additionally, portions of the vehicle  100  can be covered with a thermal protection system (TPS) to prevent structural degradation due to aerothermal heating during vehicle reentry. For example, in some embodiments the TPS can be applied to at least a fuselage nose portion  2602 , wing leading edge portions  2606   a  and  2606   b , and an underside portion  2604  of the fuselage  102  and the wings  104   a, b . Various suitable materials known in the art may be used for the TPS. By way of example, in some embodiments the TPS can include a Toughened Uni-piece Fibrous Reinforced Oxidation-resistant Composite material known as “TUFROC.” TUFROC can survive extreme heat environments up to 3,600 degrees F. and above. In other embodiments, other types of TPS materials can be used to protect vehicle  100  from aerothermal heating degradation. 
     There are a number of advantages associated with embodiments of the rocket powered launch sled  400  described above. For example, the launch sled  400  can enable the flight vehicle  100  to gain more velocity during the takeoff run than could otherwise be achieved by the unassisted vehicle  100 . Additionally, by transferring propellant to the vehicle  100  during the takeoff run, the sled  400  enables the vehicle  100  to leave the ground full (or at least nearly full) of propellant, thereby minimizing (or at least reducing) the vehicle dry weight penalty and maximizing (or at least increasing) available flight performance. In this way, the sled  400  can be viewed as a “first stage” of the vehicle  100  that never leaves the ground. The sled  400  can also provide a measure of safety during the critical first seconds of vehicle main engine firing, because the sled braking system  455  ( FIG.  4 A ) is configured to slow the vehicle  100  to a stop and/or hold it in place if anomalous performance of the main vehicle engines  120   a - c  is detected at any time during takeoff. Additionally, since the vehicle  100  takes off on the sled  400  and not its own landing gear  126   a - c , the landing gear  126   a - c  do not have to be sized to withstand the structural loads associated with takeoff with a full load of propellant. Instead, the landing gear  126   a - c  need only be sized to carry the lower loads associated with landing with empty, or near empty, propellant tanks. Accordingly, the sled  400  also eliminates the need for a heavy-duty takeoff-rated landing gear system, thereby saving flight vehicle weight and reducing system complexity. 
     The above Detailed Description of examples and embodiments of the present technology is not intended to be exhaustive or to limit the disclosed technology to the precise forms disclosed above. While specific examples for the present technology are described above for illustrative purposes, various equivalent modifications are possible within the scope of the disclosed technology, as those skilled in the relevant art will recognize. For example, while processes, routines, and/or blocks are presented in a given order, alternative implementations may perform processes and routines having steps, or employ systems having blocks, in a different order, and some processes, routines or blocks may be deleted, moved, added, subdivided, combined, and/or modified to provide alternative or sub-combinations. Each of the processes or blocks may be implemented in a variety of different ways. Also, while processes or blocks are at times shown as being performed in series, these processes or blocks may instead be performed or implemented in parallel, or may be performed at different times. 
     Several embodiments of the present technology may take the form of controller- or computer-executable instructions, such as routines executed by the vehicle controller  140  and/or the sled controller  440 , or by another data processing device, e.g., an onboard computer, special-purpose computer, server computer, personal computer, etc. Those skilled in the relevant art will appreciate that aspects of the present technology can be practiced with other computer/controller systems, including other communications, data processing, or computer system configurations. Aspects of the present technology can be embodied in a special purpose computer or data processor that is specifically programmed, configured, or constructed to perform one or more of the functions, methods, and/or computer-executable instructions explained herein. Accordingly, the terms “computer” and “controller” as generally used herein refer to any data processor and can include onboard and remote computers, 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. 
     EXAMPLES 
     The following examples provide additional embodiments of the present technology: 
     1. A method of operating a reusable space vehicle, the method comprising:
         releasably attaching the space vehicle to a launch sled;   igniting one or more rocket engines on the launch sled;   igniting one or more rocket engines on the space vehicle; and   providing propellant from one or more propellant tanks on the launch sled to the one or more rocket engines on the space vehicle and to the one or more rocket engines on the launch sled, while the one or more rocket engines on the space vehicle and the one or more rocket engines on the launch sled provide thrust for launch of the space vehicle.       

     2. The method of example 1, further comprising:
         accelerating the launch sled toward a takeoff speed of the space vehicle;   at or near the takeoff speed, ceasing to provide propellant from the one or more propellant tanks on the launch sled to the one or more rocket engines on the space vehicle; and releasing the space vehicle from the launch sled.       

     3. The method of example 2, further comprising, at or near the takeoff speed, providing propellant from one or more propellant tanks on the space vehicle to the one or more rocket engines on the space vehicle. 
     4. The method of example 2, wherein the space vehicle is configured to fly into space, orbit the Earth, and then reenter the Earth&#39;s atmosphere, and wherein the method further comprises: 
     flying the vehicle toward a landing site; 
     deploying a landing gear on the vehicle; and 
     horizontally landing the vehicle on the landing gear at the landing site. 
     5. The method of example 1, further comprising returning vented propellant from the space vehicle to the one or more propellant tanks on the launch sled as the one or more rocket engines on the space vehicle provide thrust for launch of the space vehicle. 
     6. The method of example 1 wherein the one or more propellant tanks on the launch sled includes an oxidizer tank, and wherein providing propellant from the one or more propellant tanks on the launch sled includes providing oxidizer from the oxidizer tank to the one or more rocket engines on the launch sled and to the one or more rocket engines on the space vehicle. 
     7. The method of example 1 wherein the one or more propellant tanks on the launch sled includes a fuel tank, and wherein providing propellant from the one or more propellant tanks on the launch sled includes providing fuel from the fuel tank to the one or more rocket engines on the launch sled and to the one or more rocket engines on the space vehicle. 
     8. The method of example 1 wherein the one or more propellant tanks on the launch sled include an oxidizer tank and a fuel tank, and wherein providing propellant from the one or more propellant tanks includes providing oxidizer from the oxidizer tank and fuel from the fuel tank to the one or more rocket engines on the space vehicle and to the one or more rocket engines on the launch sled. 
     9. The method of example 1 wherein the launch sled includes a first oxidizer tank and a second oxidizer tank, and wherein providing propellant from the one or more propellant tanks includes providing oxidizer from the first oxidizer tank to the one or more rocket engines on the launch sled and providing oxidizer from the second oxidizer tank to the one or more rocket engines on the space vehicle. 
     10. The method of example 1 wherein the launch sled includes a first fuel tank and a second fuel tank, and wherein providing propellant from the one or more propellant tanks includes providing fuel from the first fuel tank to the one or more rocket engines on the launch sled and providing fuel from the second fuel tank to the one or more rocket engines on the space vehicle. 
     11. The method of example 1, further comprising:
         prior to igniting the one or more rocket engines on the launch sled and the one or more rocket engines on the space vehicle, releasably attaching a propellant umbilical between the launch sled and the space vehicle, wherein providing propellant from one or more propellant tanks on the launch sled to the one or more rocket engines on the space vehicle includes flowing the propellant through the propellant umbilical;   after igniting the one or more rocket engines on the launch sled and the one or more rocket engines on the space vehicle, accelerating the launch sled toward a takeoff speed of the space vehicle;   at or near the takeoff speed, disconnecting the propellant umbilical from the space vehicle; and   releasing the space vehicle from the launch sled.       

     12. The method of example 1 wherein releasably attaching the space vehicle to the launch sled includes releasably coupling one or more support arms extending from the launch sled to the vehicle, and wherein the method further comprises: 
     supporting the space vehicle solely by the one or more support arms. 
     13. The method of example 12, further comprising rotating the one or more support arms relative to the launch sled to increase an angle of attack of the space vehicle. 
     14. The method of example 1, further comprising:
         prior to releasably attaching the space vehicle to the launch sled, rolling the vehicle onto the launch sled on a vehicle landing gear, wherein releasably attaching the space vehicle to the launch sled includes releasably coupling one or more support arms extending from the launch sled to the vehicle; and   retracting the vehicle landing gear so that the space vehicle is supported solely by the support arms.       

     15. The method of example 14, further comprising, prior to retracting the vehicle landing gear, rotating the one or more support arms relative to the launch sled to raise the vehicle off of the landing gear. 
     16. A space vehicle system, comprising:
         a reusable space vehicle having one or more rocket engines; and   a launch sled having one or more rocket engines, wherein the launch sled is configured to support the space vehicle during launch from Earth, and wherein the launch sled is further configured to provide propellant to the one or more rocket engines of the launch sled and to the one or rocket engines of the space vehicle during launch of the space vehicle.       

     17. The space vehicle system of example 16 wherein the reusable space vehicle is a horizontal takeoff/horizontal landing (HTHL) spaceplane. 
     18. The space vehicle system of example 16 wherein the reusable space vehicle is a single-stage-to-orbit (SSTO) spaceplane. 
     19. The space vehicle system of example 16 wherein the one or more rocket engines on the vehicle are bipropellant engines that use liquid oxygen and liquid fuel. 
     20. The space vehicle system of example 16:
         wherein the launch sled includes a first propellant tank,   wherein the space vehicle includes a second propellant tank that provides propellant to the one or more rocket engines on the space vehicle, and   wherein the launch sled further includes a propellant umbilical configured to transfer propellant from the first propellant tank to the second propellant tank during launch of the space vehicle.       

     21. The space vehicle system of example 20 wherein the first propellant tank provides propellant to the one or more rocket engines on the launch sled. 
     22. The space vehicle system of example 16:
         wherein the launch sled includes a first propellant tank and a second propellant tank,   wherein the first propellant tank provides propellant to the one or more rocket engines on the launch sled,   wherein the space vehicle includes a third propellant tank that provides propellant to the one or more rocket engines on the space vehicle, and   wherein the launch sled further includes a propellant umbilical configured to transfer propellant from the second propellant tank to the third propellant tank during launch of the space vehicle.       

     23. The space vehicle system of example 16 wherein the launch sled includes a first oxidizer tank, wherein the space vehicle includes a second oxidizer tank that provides oxidizer to the one or more rocket engines on the space vehicle, and wherein the launch sled is configured to provide oxidizer from the first oxidizer tank to the second oxidizer tank during launch of the space vehicle. 
     24. The space vehicle system of example 16 wherein the launch sled includes a first fuel tank, wherein the space vehicle includes a second fuel tank that provides fuel to the one or more rocket engines on the space vehicle, and wherein the launch sled is configured to provide fuel from the first fuel tank to the second fuel tank during launch of the space vehicle. 
     25. The space vehicle system of example 16:
         wherein the launch sled includes one or more support arms configured to releasably support the space vehicle on the launch sled,   wherein a distal end portion of each of the one or more support arms includes a hold and release mechanism configured to releasably attach to a corresponding interface on the space vehicle, and   wherein each of the one or more support arms is movable to change the position of the space vehicle relative to the launch sled.       

     26. The space vehicle system of example 16:
         wherein the launch sled includes a first propellant tank,   wherein the space vehicle includes a second propellant tank that provides propellant to the one or more rocket engines on the space vehicle,   wherein the launch sled further includes a propellant umbilical configured to transfer propellant from the first propellant tank to the second propellant tank during launch of the space vehicle, the propellant umbilical extending between a propellant outlet interface on an upper portion of the launch sled to a propellant inlet interface on an underside of the space vehicle,   wherein the launch sled further includes at least one support arm having a proximal end portion pivotally coupled to the upper portion of the launch sled laterally adjacent to the propellant outlet interface, and a distal end portion configured to be releasably coupled to an interface fitting on the underside of the space vehicle laterally adjacent to the propellant inlet interface, and   wherein the at least one support arm is rotatable to change the position of the space vehicle relative to the launch sled.       

     27. A reusable space vehicle, comprising:
         a pair of wings configured to provide lift during flight of the space vehicle in the Earth&#39;s atmosphere;   one or more rocket engines; and   an oxidizer tank configured to provide liquid oxygen to the one or more rocket engines, the oxidizer tank having a non-cylindrical and non-spherical shape.       

     28. The reusable space vehicle of example 27, further comprising a fuselage having an external sidewall, and wherein the external sidewall forms a portion of the oxidizer tank. 
     29. The reusable space vehicle of example 27 wherein the oxidizer tank is configured to withstand an internal pressure of less than 4 psig. 
     30. The reusable space vehicle of example 27 wherein the oxidizer tank is configured to carry liquid oxygen at a temperature of from about −196 degrees C. to about −182 degrees C. during launch of the space vehicle. 
     31. The reusable space vehicle of example 27, further comprising one or more structural interfaces on an underside thereof configured to releasably attach the space vehicle to a rocket-powered launch sled for takeoff of the space vehicle. 
     32. The reusable space vehicle of example 27, further comprising:
         one or more structural interfaces on an underside thereof configured to releasably attach the space vehicle to a rocket-powered launch sled for takeoff of the space vehicle; and   a landing gear that is only used for landing the space vehicle.       

     33. A launch sled for launching a reusable space vehicle, the space vehicle having one or more rocket engines, the launch sled comprising:
         one or more rocket engines for providing thrust to the launch sled during launch of the vehicle; and   a propellant tank configured to be operably coupled in fluid communication with the space vehicle, wherein the propellant tank is configured to provide propellant to the one or more engines of the space vehicle during launch of the space vehicle.       

     34. The launch sled of example 33, further comprising a propellant umbilical configured to be releasably connected between the launch sled and the space vehicle to transfer propellant therebetween during launch of the space vehicle. 
     35. The launch sled of example 33, further comprising a plurality of movable support arms configured to releasably attach the space vehicle to the launch sled and move the space vehicle relative to the launch sled. 
     36. A method for loading liquid oxygen into a horizontal takeoff/horizontal landing space vehicle, the method comprising:
         cooling the liquid oxygen to a temperature of about −196 degrees C. or less; and   flowing the cooled liquid oxygen into an oxidizer tank on the space vehicle.       

     37. The method of example 36 wherein cooling the liquid oxygen includes flowing the liquid oxygen through a liquid nitrogen heat exchanger, and wherein the method further comprises:
         venting oxygen from the oxidizer tank;   recirculating the vented oxygen through the liquid nitrogen heat exchanger to re-cool the liquid oxygen to a temperature of about −196 degrees C. or less; and   flowing the re-cooled liquid oxygen back to the oxidizer tank.       

     38. The method of example 36 wherein the space vehicle is mounted to a launch sled having a first oxidizer tank, wherein the oxidizer tank on the space vehicle is a second oxidizer tank, and wherein the method further comprises flowing the cooled liquid oxygen into the first oxidizer tank on the sled and then flowing the cooled liquid oxygen from the first oxidizer tank into the second oxidizer tank. 
     39. The method of example 36 wherein flowing the liquid oxygen into the oxidizer tank includes flowing the liquid oxygen into an oxidizer tank having a non-cylindrical, non-spherical shape. 
     40. The method of example 36 wherein flowing the liquid oxygen into the oxidizer tank includes flowing the liquid oxygen into an oxidizer tank on a space vehicle having a pair of wings for horizontal takeoff and landing in the Earth&#39;s atmosphere. 
     41. The method of example 36 wherein flowing the liquid oxygen into the oxidizer tank includes flowing the liquid oxygen into an oxidizer tank in a fuselage of a space vehicle having a pair of wings for horizontal takeoff and landing in the Earth&#39;s atmosphere, the oxidizer tank having a non-cylindrical, non-spherical shape. 
     References throughout the foregoing description to features, advantages, or similar language do not imply that all of the features and advantages that may be realized with the present technology should be or are in any single embodiment. Rather, language referring to the features and advantages is understood to mean that a specific feature, advantage, or characteristic described in connection with an embodiment is included in at least one embodiment of the present technology. Thus, discussion of the features and advantages, and similar language, throughout this specification may, but do not necessarily, refer to the same embodiment. Furthermore, the described features, advantages, and characteristics of the present technology may be combined in any suitable manner in one or more embodiments. One skilled in the relevant art will recognize that the present technology can be practiced without one or more of the specific features or advantages of a particular embodiment. In other instances, additional features and advantages may be recognized in certain embodiments that may not be present in all embodiments of the present technology. 
     Any patents and applications and other references noted herein, including any that may be listed in accompanying filing papers, are incorporated herein by reference. To the extent that any materials incorporated herein by reference conflict with the present disclosure, the present disclosure controls. Aspects of the present technology can be modified, if necessary, to employ the systems, functions, and concepts of the various references described above to provide yet further implementations of the present technology. 
     While the above description describes various embodiments of the disclosed technology and the best mode contemplated, regardless how detailed the above text, the technology can be practiced in many ways. Details of the system may vary considerably in its specific implementation, while still being encompassed by the present disclosure. As noted above, particular terminology used when describing certain features or aspects of the disclosed technology should not be taken to imply that the terminology is being redefined herein to be restricted to any specific characteristics, features, or aspects of the disclosed technology with which that terminology is associated. In general, the terms used should not be construed to limit the disclosed technology to the specific examples disclosed in the specification, unless the above Detailed Description section explicitly defines such terms. As used herein, the term “and/or”, as in “A and/or B,” refers to A alone, B alone, and both A and B. 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 spirit and scope of the present disclosure. 
     From the foregoing, it will be appreciated that specific embodiments of the invention have been described herein for purposes of illustration, but that various modifications may be made without deviating from the spirit and scope of the various embodiments of the invention. Further, while various advantages associated with certain embodiments of the invention have been described above 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 invention. Accordingly, the invention is not limited, except as by the appended claims. 
     Although certain aspects of the invention are presented below in certain claim forms, the applicant contemplates the various aspects of the invention in any number of claim forms. Accordingly, the applicant reserves the right to pursue additional claims after filing this application to pursue such additional claim forms, in either this application or in a continuing application.