Abstract:
An aerospace vehicle for delivering a payload into space includes a first stage and a second stage with a plurality of separation assemblies coupling the first stage to the second stage. At least one container charged with pressurized gas in fluid communication with the separation assemblies provides pressurized gas to the separation assemblies to cause separation of the first stage and the second stage.

Description:
TECHNICAL FIELD OF THE INVENTION 
     The present invention relates to aerospace vehicles and, in particular, to a stage separation system and method. 
     BACKGROUND OF THE INVENTION 
     Multistage aerospace vehicles are widely used to carry payloads into orbit and propel space vehicles into outer space. One or more booster stages accelerate an orbital stage, or vehicle, toward space. Depleted booster stages are generally dropped in order to reduce the weight of the aerospace vehicle. After each booster stage has served its purpose in attaining a certain velocity, it separates from the next stage and falls back to earth. 
     Timely and proper separation of stages in an aerospace vehicle often requires intricate planning and design, and typically involves high-cost, sensitive hardware and instrumentation. Separation is often accomplished by detonating pyrotechnic devices in a predetermined sequence which in turn disengage the mechanical connection between stages. 
     Pyrotechnic devices, however, are hazardous explosives, and inherently expensive to manufacture, deliver and handle. Therefore, the number of pyrotechnic devices employed in a given system has significant cost implications. Furthermore, the shock and debris of pyrotechnic devices may have a deleterious effect on other system components including the booster stage(s) and orbital vehicle because they cause structural damage above and beyond that required for separation. This collateral damage increases with the number of devices utilized and impacts the ability to reuse system components for subsequent launches. 
     SUMMARY OF THE INVENTION 
     An object of the present invention is to reduce the cost of placing aerospace vehicles or payloads in earth orbits and space, and in particular, to provide a multistage separation system which employs a limited number of pyrotechnic devices. Another object is to enhance the efficiency of separation of multistage aerospace vehicles. Yet another object is to minimize damage to aerospace vehicles caused by pyrotechnic devices. Still another object is to provide a safe, reliable, cost-effective separation system for multistage aerospace vehicles. 
     The foregoing objects are attained in accordance with the present invention by employing a separation system which requires a limited number of pyrotechnic devices. In a particular embodiment, a first stage and a second stage are provided with a plurality of separation assemblies coupling the first stage to the second stage. At least one container charged with pressurized gas in fluid communication with the separation assemblies may also be provided. The container provides pressurized gas to the separation assemblies to cause separation of the first stage and the second stage. 
     In another embodiment of the present invention, a first stage and a second stage may be coupled to define a cavity between the first stage and the second stage. An orifice operable to provide fluid communication between the cavity and external ambient may also be provided. 
     In yet another embodiment, a separation system for use on a multistage aerospace vehicle includes a manifold and a plurality of containers filled with pressurized gas, in fluid communication with the manifold. A plurality of separation nut assemblies in fluid communication with the manifold are also provided. A plurality of valves are disposed between the manifold and the containers. The valves, upon actuation, release the pressurized gas for delivery to the separation nut assemblies. 
     A technical advantage of the present invention includes the limited number of pyrotechnic devices required to effectuate the safe and efficient separation of the stages. By limiting the number of pyrotechnic devices, collateral damage to the structural components of the aerospace vehicle is minimized, as well as the amount of flying debris generated. This allows the operator to refurbish and reuse the aerospace vehicle during subsequent launches. 
     Another technical advantage includes separation of the stages using trapped air to prevent unwanted side velocities, uneven separation, and structural damage. The trapped air may be controllably released to establish predetermined separation forces. 
     Other technical advantages are readily apparent to one skilled in the art from the following figures, description, and claims. 
    
    
     BRIEF DESCRIPTION OF THE DRAWINGS 
     For a more complete understanding of the present invention, and for further features and advantages, reference is now made to the following description, taken in conjunction with the accompanying drawings, in which: 
     FIG. 1 is a side elevation view of a launch vehicle that includes an orbital vehicle and launch assist platform embodying the present invention; 
     FIG. 2 is a partial cross-section, with portions broken away, illustrating a portion of the juncture between the orbital vehicle and launch assist platform; 
     FIG. 3 is a generally schematic view of the juncture between the orbital vehicle and launch assist platform taken along lines  3 — 3  of FIG. 1; 
     FIG. 4 is a partial perspective view from a point on the interior of the launch assist platform; 
     FIG. 5 is an electrical/pneumatic interconnect block diagram; 
     FIG. 6 is a cross-sectional view of a “D” seal, in an undeformed state; 
     FIG. 7 is an exploded perspective view, with portions broken away, illustrating pressure control orifices; and 
     FIG. 8 is a partial perspective view, with portions broken away, looking from a point to the side and above the launch assist platform, illustrating a separation nut assembly accessible from the outside of the vehicle. 
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     Referring to FIG. 1, a launch vehicle  30  is illustrated which includes an orbital vehicle  32  and a booster stage, or launch assist platform  34 , which propels orbital vehicle  32  toward an orbit around the earth. The juncture of orbital vehicle  32  and launch assist platform  34  is indicated by cross-section  3 — 3  and further illustrated in FIG.  3 . The two-stage combination of orbital vehicle  32  and launch assist platform  34  delivers a payload  33  into earth orbit. Launch assist platform  34  may be used alone or in combination with one or more additional booster stages to assist a space vehicle in reaching earth orbit or outer space. While the illustrated embodiment encompasses a two-stage launch vehicle, the teachings of the present invention apply to separation techniques and structures between stages of an aerospace vehicle. 
     In the present embodiment, launch assist platform  34  includes a body that is essentially a tubular, aerodynamic outer shell  36  of cylindrical shape which is constructed in major part by internal ribbed tubular panels of a composite material. Launch assist platform  34  derives its power from one or more liquid oxygen/kerosene main propulsion engines  38 . A liquid oxygen (LOX) propellant tank  41  and kerosene fuel tank  40  are encased within launch assist platform  34 , and thermally isolated from the ribbed tubular panels of outer shell  36 . Liquid oxygen stored within propellant tank  41  and kerosene stored within fuel tank  40  is supplied to main propulsion engines  38  to provide thrust to launch assist platform  34  during take-off and flight of launch vehicle  30 . 
     The propulsion system  42  associated with orbital vehicle  32  includes engine  44 , liquid oxygen propellant tank  47  and kerosene fuel tank  46 . After separation of launch assist platform  34  from orbital vehicle  32 , liquid oxygen stored within propellant tank  47  and kerosene stored within fuel tank  46  are supplied to engines  44  to provide thrust to orbital vehicle  32 . The orbital vehicle includes a blunt nose  50 , which is generally parabolic shaped, and outer shell  48  formed from ribbed tubular panels of a composite material. 
     Orbital vehicle  32  and launch assist platform  34  define a cavity  110  at their juncture. An orifice  118  provides fluid communication between cavity  110  and external ambient. Air and/or pressurized gas trapped within cavity  110  may be selectively released through orifice  118 . Pressure relief valves  128  may be used in lieu of, or in addition to orifice  118 , to maintain a specified pressure differential between cavity  110  and external ambient. This trapped air pressure release system will be described later, in more detail. 
     In one embodiment, launch vehicle  30  may be used to deliver communications satellites into low earth orbit. The components of launch vehicle  30  may be fully reusable in excess of one hundred launch applications. Launch vehicle  30  of FIG. 1 is approximately one hundred fifteen feet in overall length, twenty-two feet in diameter and may weigh in excess of eight hundred and five thousand pounds at lift off. During operation, main engines  38  provide the thrust necessary to achieve lift off and sustain flight of launch vehicle  30  to a predetermined elevation and trajectory. A separation system to be described in more detail later decouples orbital vehicle  32  from launch assist platform  34 . Main engines  38  provide the necessary thrust to maneuver launch assist platform  34  to a predetermined location where a chute and airbag system is deployed, which allows launch assist platform  34  to safely return to the earth&#39;s surface for recovery and reuse. 
     Shortly after separation, engine  44  ignites to propel orbital vehicle  32  into earth orbit. Payload  33  is then deployed to remain in an orbital trajectory. A de-orbit burn provided by an orbital maneuvering engine then allows orbital vehicle  32  to exit earth&#39;s orbit and return to earth. At a predetermined elevation, another chute and airbag system deploys to allow orbital vehicle  32  to land safely on the earth&#39;s surface. Orbital vehicle  32  and launch assist platform  34  are then collected to be retrofitted and refueled for another launch sequence to deploy an additional payload. 
     FIG. 2 illustrates a portion of the juncture between launch assist platform  34  and orbital vehicle  32 . A flanged portion  60  of launch assist platform  34  couples to a flanged portion  62  of orbital vehicle  32  using separation nut assembly  70 . Separation nut assembly  70  includes an outer housing  72  surrounding a retainer spring  74 . A cartridge port  78  connects to branch piping section  102 . Pressure chamber  80  provides a fluid communication path between branch piping section  102  and separation bolt  76 . 
     A bolt retainer  90  is optionally provided and mounted to orbital vehicle  32  to capture and retain separation bolts  76  upon actuation of separation nut assembly  70  and separation of launch assist platform  34  from orbital vehicle  32 . Actuation of separation nut assembly  70  releases bolt  76  which is retained within bolt retainer  90 . This helps minimize the amount of flying debris generated during the separation stage of the launch which could otherwise damage structural components of launch vehicle  30  and create hazards for other aircraft, as well as structures and populations below. Separation nut assemblies  70  are actuated by introducing pressurized gas through branch piping section  102  to cartridge port  78  of separation nut assembly  70 . 
     FIG. 3 is a cross section and illustrates portions of separation system  101 . Manifold  100  is installed along the perimeter of outer shell  36  of launch assist platform  34  near the interface between launch assist platform  34  and orbital vehicle  32 . Manifold  100  includes a circular tube with a hollow generally tubular cross-section. A plurality of branch piping sections  102  provide fluid communication paths between manifold  100  and separation nut assemblies  70 . Pressurized containers  106  secure to the interior of outer shell  36  of launch assist platform  34 . In one embodiment, pressurized tank  106  is charged with nitrogen gas (N 2 ), but it should be recognized that other gases, including helium gas (H 2 ), can be utilized. Branch piping sections  108  provide fluid communication paths between pressurized containers  106  and manifold  100 . 
     Pyrotechnic valves  104  are disposed within branch piping sections  108  and maintained in a typically “closed” position until actuation of the separation nut assemblies  70  is desired. A PCR 1/2-20 Power Cartridge as produced by Hi-Shear Technology Corporation, for example, is suitable for use within the teachings of the present invention. Upon actuation of pyrotechnic valves  104 , fluid communication is established between pressurized containers  106  and manifold  100 , allowing trapped gas to travel through pyrotechnic valves  104  which are in the “open” position, through branch piping sections  108  fully charging manifold  100  almost instantaneously. Pressurized gas then proceeds through branch piping sections  102  into separation nut assembly  70  via cartridge port  78 . In another embodiment, mechanical and/or electromechanical valves may be used interchangeably with, or instead of pyrotechnic valves  104 . 
     In one embodiment, pressurized containers  106  may be charged with Helium gas to a pressure of 7,500 psi. Nitrogen gas may also be introduced into manifold  100  prior to launch, to affect a faster overall charge. As an example, manifold  100  may be pre-charged with nitrogen gas to a pressure of 2,500 psi. When helium gas within pressurized container  106  and nitrogen gas within manifold  100  are used in combination, the chemical reaction caused by the mixing of the gases enhances the performance of the system facilitating more rapid actuation of separation nut assemblies  70 . It will be recognized by those of ordinary skill in the art that many types of compressed gas are available for use interchangeably within manifold  100  and pressurized containers  106 . 
     FIG. 4 illustrates a partial perspective view of a portion of separation system  101 . Although the fluid communication path described includes manifold  100 , and branch piping sections  102  and  108 , it should be recognized by those of ordinary skill in the art that any reference to a manifold may include any piping, fittings, and branch lines necessary to allow fluid communication between pressurized containers  106  and separation nut assemblies  70 . 
     FIG. 5 illustrates a piping and instrumentation diagram of separation system  101 . A battery pack  82  provides power to a controller or central processing unit (CPU)  84  which controls the actuation of pyrotechnic valves  104 . Upon command, CPU  84  actuates pyrotechnic valves  104  using signal lines  103 , allowing gas contained within pressurized containers  106  to enter branch piping sections  102 , charging manifold  100 . Separation nut assemblies release when a predetermined amount of pressure is transferred from manifold  100  through branch piping sections  102  to separation nut assemblies  70 . CPU also includes redundant sensor lines  105  to monitor the pressure of compressed gas in pressurized containers  106 . Any number of specific configurations of the components of separation system  101  are available in lieu of the system illustrated in FIG.  5 . As an example, valves may be provided within manifold  100  essentially partitioning the system such that each pressurized container  106  services a specific number of separation nut assemblies. In one embodiment, four pressurized containers may be employed to service a total of twenty-four separation nut assemblies which would allow a design wherein each pressurized container services a total of six separation nut assemblies. In another embodiment, the ratio of separation assemblies to containers charged with pressurized gas may exceed 6:1. In order to avoid errors or complications caused by faulty components, redundancy may also be introduced into the separation system components. For example, each pressurized container may service six primary separation nut assemblies  70  and also provide “backup” to an additional six in case of equipment failure. 
     In one particular embodiment, low shock separation nuts within the SN9400 Series as manufactured by Hi-Shear Technology Corporation of Torrance, Calif. are suitable for as separation assemblies in separation system  101 . Such bolts facilitate a torque of 140 foot-pounds applied to the mechanical connection between orbital vehicle  32  and launch assist platform  34 . In another embodiment, separation assemblies  70  may be programmed to release when pressure in the range of four to five thousand pounds per square inch is introduced at the cartridge port. In a particular embodiment, release of all separation assemblies  70  associated with separation system  101  may then be accomplished in less than eight milliseconds. Many releasable mechanical couplings, or separation assemblies, are available for use as separation assemblies, within the teachings of the present invention. 
     Once the structural bond of separation assemblies  70  is broken, the physical separation of orbital vehicle  32  from launch assist platform  34  of launch vehicle  30  of FIG. 1 is enhanced by a volume of trapped air occupying interior cavity  110  defined by components of launch assist platform  34  and orbital vehicle  32 . This volume of air may be maintained at a predetermined pressure. Throughout the flight of launch vehicle  30 , the pressure within interior cavity  110  remains higher than ambient atmospheric pressure since ambient atmospheric pressure will decrease steadily corresponding to any increase in elevation. The pressure within interior cavity  110  may be controlled passively through orifice  118  or pressure relief valves  128  or actively using sensors. 
     Launch vehicle  30  includes interior cavity  110  (FIG. 1) which occupies the space between and within portions of orbital vehicle  32  and launch assist platform  34 . Interior cavity  110  is defined at its perimeter by outer shells  36  and  48  of launch assist platform  34  and orbital vehicle  32 , respectively. The lower boundary of interior cavity  110  is defined by propellant tank  41  of launch assist platform  34  and the upper extreme is defined by fuel tank  46  of orbital vehicle  32 . Launch assist platform  34  is assembled in a manner in which air cannot pass between propellant tank  41  and outer shell  36 . Similarly, orbital vehicle  32  is assembled such that air is prevented from traveling between fuel tank  46  and outer shell  48 . Although many components of launch vehicle  30  occupy interior cavity  110 , including propulsion system  42  of orbital vehicle  32  and other components, a significant volume remains wherein air and other gases may be contained. 
     When launch vehicle  30  is assembled prior to launch, the juncture between launch assist platform  34  and orbital vehicle  32  forms a generally airtight seal. A circular notched opening  112  (FIG. 2) with a generally rectangular cross section is provided near the outermost perimeter of flanged portion  60  of launch assist platform  34 . A similar circular notched opening  114  (FIG. 2) is provided near the innermost perimeter of flanged portion  60  of launch assist platform  34 . A pair of circular “D” seals  116 , the cross-section of which is illustrated in FIG. 6, are inserted into circular notched openings  112  and  114  during the assembly of launch vehicle  30 . When separation bolt  76  is torqued down, flanged portion  62  of orbital vehicle  32  compresses “D” seals  116  within circular notched openings  112  and  114 , thereby creating a generally airtight seal between flanged portion  62  of orbital vehicle  32  and flanged portion  60  of launch assist platform  34 . Although the illustrated embodiment utilizes two “D” seals  116  to close any opening which may exist between flanged portion  60  and flanged portion  62 , a single “D” seal may be sufficient. Alternatively, it will be recognized by those skilled in the art that many other methods of establishing this generally airtight seal are available. For example, flanged portions  60  and  62  may be machined in such a manner that “D” seal  116  would not be required to establish a substantially airtight seal. 
     In one embodiment, circular notched openings  112  and  114  may have cross-sectional dimensions of 0.312″ wide by 0.3″ tall. Within the same embodiment, “D” seals  116  may have a cross-sectional overall width of 0.31″ and overall height of 0.5″. Separation system  101  may incorporate any number, shape, size and configuration of circular notched openings  112  and  114 , as well as “D” seals  116 , to provide a substantially airtight seal. “D” seal  116  of the illustrated embodiment is suitable to fill manufacturing, assembly, and frame gapping of approximately 0.121″. 
     After the assembly of launch vehicle  30  prior to launch, interior cavity  110  is substantially airtight with respect to ambient atmospheric pressure. Accordingly, the pressure within interior cavity  110  will remain at whatever ambient pressure is prevalent at the elevation where assembly is accomplished. This pressure may fall within the range of 10-15 psi according to the assembly and launch sites currently being contemplated. Once interior cavity  110  is sealed, this pressure may be maintained regardless of changes encountered in ambient atmospheric pressure due to changes in elevation experienced during the launch and flight of launch vehicle  30 . 
     In order to selectively control the dissipation of pressure within cavity  110 , an orifice  118  may be provided within outer shell  36  of launch assist platform  34 . Orifice  118  provides a fluid communication path between interior cavity  110  and the ambient atmosphere. Orifice  118  may be located anywhere along the perimeter of interior cavity  110  along either outer shell  36  of launch assist platform  34  or outer shell  48  of orbital vehicle  32  or both. Any number of the same or differently sized additional orifices may also be employed, although the illustrated embodiment contemplates the use of a single orifice  118 . The appropriate size of orifice  118  will depend upon a number of factors including, but not limited to, its location upon launch vehicle  30 , the elevation of the assembly and launch, the elevation at which the separation will be accomplished, the time from launch to separation, the amount of pressure necessary to accomplish the physical separation of orbital vehicle  32  from launch assist platform  34 , the effectiveness of the generally airtight seal for cavity  110 , and other fluid dynamic characteristics associated with the launch and flight of launch vehicle  30 . The use of orifice  118  is not required to affect the separation of orbital vehicle  32  from launch assist platform  34 , but provides a mechanism by which pressure within interior cavity  110  may be passively controlled to a predetermined level at separation. 
     FIG. 7 illustrates a plate  120  of a composite or metallic material. Since launch vehicle  30  is intended to be fully reusable and the fluid dynamics associated with each flight may vary significantly, the illustrated embodiment facilitates rapid modification and interchangeability of the size of orifice  118 . Plate  120  is suitable for installation upon launch assist platform  34 . In order to allow pressure dissipation from within interior cavity  110 , a fixed orifice  122  is provided within launch assist platform  34 . Composite plate  120  is then installed over fixed orifice  122  such that composite plate  120  completely covers fixed orifice  122 . A variable orifice  124  is provided within composite plate  120  and aligned with fixed orifice  122  to control the dissipation of pressure from within interior cavity  110 . Fixed orifice  122  may be provided at any size suitable to be completely covered by composite plate  120 . Variable orifice  124  controls pressure dissipation from interior cavity  110  and its size is therefore controlling in the design of the required pressure dissipation system. 
     Variable orifice  124  is provided within composite plate  120  to accommodate the rapid interchangeability of various variable orifice sizes. When a different size variable orifice is required due to specific design considerations, composite plate  120  may be removed from launch assist platform  34  quickly and efficiently by removing mechanical fasteners  126 . Another composite plate  220  with a different size variable orifice  224  may then be installed upon launch assist platform  34 . 
     For another launch with different launch conditions, fixed orifice  122  and therefore, composite plate  120  containing variable orifice  124 , may be installed anywhere within the perimeter of interior cavity  110  provided fluid communication with an area of lower pressure is provided. The position of any orifice may be adjusted due to the dynamics of hypersonic flows and vortexing. In the illustrated embodiment, composite plate  120  is provided along the upper perimeter of outer shell  48  of launch assist platform  34  by way of example only. 
     Another method for selectively controlling pressure dissipation from within interior cavity  110  uses one or more pressure relief valves  128  (FIG. 1) installed upon the outer perimeter of interior cavity  110 . Pressure relief valves  128  form a fluid communication path between interior cavity  110  and the ambient atmosphere. Pressure relief valves  128  are preset to allow pressure within interior cavity  110  to escape to the atmosphere until a desired pressure differential across pressure relief valve  128  is accomplished. In this manner, the pressure differential between interior cavity  110  and ambient atmosphere can be maintained at a predetermined level to ensure the optimum performance of the trapped air pressure separation system. 
     As an example, pressure relief valves  128  may be preset to maintain a pressure differential of approximately 3 to 8 psi, ensuring that the pressure within interior cavity  110  will remain 3 to 8 psi higher than ambient atmospheric pressure at all times during flight. In this manner, the volume of trapped air within interior cavity  110  between orbital vehicle  32  and launch assist platform  34  is allowed to retain pressure 6.5 psi greater than ambient and this pressure is used to force the stages apart upon separation. In another embodiment of the present invention, cavity  110  may be pre-charged with air or gas to maintain a higher pressure than ambient launch pressure. The shape and configuration of engines  44  further enhance the separation of stages from a “plunger” type effect which forces gasses out around the nozzle of engine  44 , upon separation. In one embodiment, a distance of 150′ to 200′ may be achieved between orbital vehicle  32  and launch assist platform  34  prior to ignition of engine  44 . 
     Although the illustrated embodiment includes one pressure relief valve, the number, size, specifications and location of the pressure relief valves may be significantly modified to achieve various design goals for a particular launch and flight. For some applications, no pressure relief valves are required. Furthermore, many other methods are available for controlling the pressure differential between ambient atmospheric pressure and the pressure within interior cavity  110 . For a more active control, a pressure transducer  130  (see FIG. 1) may be installed within interior cavity  110  in order to determine the pressure within cavity  110 . A control valve may also be provided in lieu of pressure relief valve  128  to maintain or decrease the pressure within interior cavity  110 , in response to signals from pressure transducer  130 . 
     As illustrated in FIG. 8, launch vehicle  30  may be modified to allow for convenient and rapid adjustment of separation assemblies  70  by allowing access from the exterior of launch vehicle  30 . As illustrated in FIG. 8, outer shell  36  of launch assist platform  34  may be provided with a rectangularly shaped recess  52  around each separation assembly  70 . Final adjustment and torque of separation assemblies  70  may then be accomplished after assembly of launch vehicle  30 . Furthermore, separation assemblies may be provided which allow for access through the separation assembly to the threaded end of separation bolt  76  for preloading. Accordingly, the time required for assembly and/or disassembly is drastically reduced. Recess  52  may also be utilized to provide access to install, remove and/or replace pyrotechnic valves  104  without disassembling orbital vehicle  32 . 
     Although the present invention has been described in several embodiments, a myriad of changes, variations, alterations, transformations, and modifications may be suggested to one skilled in the art, and it is intended that the present invention encompass such changes, variations, alterations, transformations, and modifications as fall within the spirit and scope of the appended claims.