WO2003086860A9 - Commercial space transportation system - Google Patents

Commercial space transportation system

Info

Publication number
WO2003086860A9
WO2003086860A9 PCT/GB2003/001582 GB0301582W WO03086860A9 WO 2003086860 A9 WO2003086860 A9 WO 2003086860A9 GB 0301582 W GB0301582 W GB 0301582W WO 03086860 A9 WO03086860 A9 WO 03086860A9
Authority
WO
WIPO (PCT)
Prior art keywords
aerospaceplane
lox
present
stage
oxidizer
Prior art date
Application number
PCT/GB2003/001582
Other languages
French (fr)
Other versions
WO2003086860A1 (en
Inventor
Jason Andrews
Dana Andrews
Original Assignee
Andrews Space And Technology I
Miller James Lionel Woolverton
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Andrews Space And Technology I, Miller James Lionel Woolverton filed Critical Andrews Space And Technology I
Publication of WO2003086860A1 publication Critical patent/WO2003086860A1/en
Publication of WO2003086860A9 publication Critical patent/WO2003086860A9/en

Links

Classifications

    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/14—Space shuttles
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/002—Launch systems
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
    • B64G1/40—Arrangements or adaptations of propulsion systems
    • B64G1/401—Liquid propellant rocket engines
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
    • B64G1/40—Arrangements or adaptations of propulsion systems
    • B64G1/402—Propellant tanks; Feeding propellants

Definitions

  • the present invention relates to the field of aerospace vehicles, and more particularly, to a novel method and apparatus of placing a vehicle in low-earth orbit.
  • rockets can be used to launch space vehicles and to place satellites into earth orbit .
  • a propellant mixes with an oxidizer and burns to provide thrust. All fuel must combine with an oxidizer in order to burn. At sea level, a separate oxidizer is not needed, as ambient air can serve as the oxidizer. At high altitudes (and in space) , however, the air thins and there is no readily available oxidizer which allows the fuel to burn. It is therefore necessary for the spacecraft to carry its own oxidizer.
  • LH 2 liquid hydrogen
  • LOX liquid oxygen
  • oxidizer can be manufactured from ambient air as a spacecraft is flying. This prior art method, however, can only be used at high supersonic speeds. Flying a craft at supersonic speeds is generally not desirable.
  • the present invention overcomes the limitations of the prior art by providing a method and apparatus which permits a space vehicle to take off without an initial load of oxidizer.
  • the spacecraft cruises at a predetermined altitude at subsonic speed.
  • the spacecraft carries an onboard liquid oxygen (LOX) generator.
  • the LOX generator uses liquid hydrogen fuel to cool incoming air, and liquefy the gaseous oxygen. This liquid oxygen is then stored in a tank until it is needed to serve as an oxidizer in a rocket engine.
  • the present invention provides a novel space transportation architecture which overcomes the limitations of the prior art .
  • the architecture includes an areospaceplane capable of horizontal take-off and landing.
  • the areospaceplane includes a device for generating liquid oxygen from ambient air while in flight . Either an orbiter or an expendable upper stage can be coupled to the areospaceplane.
  • Figure 1 is a perspective view showing an aerospaceplane of the present invention.
  • Figure 2 is a perspective view showing an orbiter 12 of the present invention.
  • Figure 3 is a perspective view showing a cargo upper stage 14 of the present invention.
  • Figure 4 is a diagram of a typical mission scenario of a spacecraft employing the method of the present invention.
  • Figure 5 is a table which lists the major engine specifications used in the present invention.
  • Figure 6 illustrates an idealized view of the booster 16 used in the system of the present invention.
  • Figure 7 is a plan view of the booster 16.
  • Figure 8 illustrates an idealized view of a second stage orbiter 18 used in the present invention.
  • Figure 9 is a plan view and an elevation of the orbiter 18.
  • Figure 10 illustrates an expendable upper stage 20 which can be used in place of the orbiter.
  • Figure 11 shows the structural design of the upper stage 20.
  • the present invention allows a reusable space vehicle to attain low-earth orbit in a manner which requires the use of a minimum amount of fuel for a given amount of payload.
  • the preferred embodiment of the present invention includes an areospaceplane as is shown in Figure 1.
  • the areospaceplane is able to take off and land in a manner like a conventional airplane .
  • the spacecraft therefore includes wings and horizontal stabilizers and a vertical stabilizer.
  • the areospaceplane is a piloted vehicle, equipped with air breathing engines.
  • the areospaceplane also includes a liquid oxygen (LOX) generation system.
  • LOX liquid oxygen
  • One such LOX generation system is described in United States Patent Application Serial No. 09/515,951, which is the preferred embodiment of the invention. It will be understood that other methods of generating LOX can be used with equal ef ectiveness. It is anticipated that the present invention can utilize wide variety of different mechanical means to compress and cool the air to form LOX.
  • Figure 6 shows an idealized form of the booster.
  • the booster is capable of a total of ⁇ v of
  • Rocket engine performance is expected to degrade by about 10% from the nominal specifications, due to the adjustment in oxidizer chemistry required by the LOX generation and propulsion system.
  • the booster's airframe in the preferred embodiment is a combination of aluminum and composite structural elements .
  • the total wing in the preferred embodiment is 511 m 2 (5500 sq. ft) , but it will be apparent to those of skill in the art that a larger or smaller wing planform can be used. Due to the LOX tank being empty at take-off as well as landing, the vehicle is capable of utilizing a standard Boeing 777 landing gear.
  • Figure 7 shows a plan view of the preferred design for the booster.
  • the vehicle is equipped with four turbo-fan engines, modified to operate on standard jet fuel as well as the gaseous hydrogen fuel provided by the LOX generation system.
  • the system's second stage orbiter is shown in Figure 8.
  • the second stage is a pressurized orbital vehicle, capable of transporting cargo as well as human crew.
  • the preferred design uses two rocket - engines (which, in the preferred embodiment are identical to those on the booster) for its main propulsion system. This gives the orbiter a total anticipated ⁇ v capability of 5.5 km/s (18,150 fps) .
  • OMS Orbital Maneuvering System
  • ACS Attitude Control System
  • the airfra e of the orbiter is constructed entirely of composites, including an integral forward crew compartment capable of transporting up to 8,165 kg (18,000 lbs.) to and from orbit.
  • the wing surface area in the preferred embodiment is approximately 93 m 2 (1000 sq. ft) .
  • Figure 9 shows a top and side view of the design.
  • Figure 10 illustrates an expendable upper stage which can be used in place of the orbiter of Figure 8.
  • the upper stage is an unpressurized orbital vehicle intended for transporting large and heavy payloads, such as GEO satellites and ISS-class habitation modules. It can be used in a reusable configuration utilizing a deployable recovery system, or as a completely. second stage in order to increase the delivered payload capability.
  • the upper state uses two rocket engines similar to those in the booster, together with a non- oxic RCS. It does not include a separate OMS in its current design.
  • the structural design of the Upper Stage is shown in Figure 11. It incorporates a 18.3 (60 ft) payload fairing, capable of supporting up to 27,215 kg (60,000 lbs.).
  • the orbiter is first mated to the areospaceplane.
  • the areospaceplane then takes off from a traditional airport facility, and climbs to a cruising altitude of approximately 25K feet.
  • the LOX system is then used to generate fuel for the rocket engine.
  • the rocket engine is used to lift the areospaceplane to altitude.
  • the vehicles separate.
  • the orbiter continues to climb until it reaches the desired low earth orbit .
  • the areospaceplane reenters the atmosphere and lands at the nearest airport .
  • FIG. 12 A schematic diagram of the mission profile is shown in Figure 12.
  • the system takes off with a compliment of LH 2 and JP-8 and climbs using conventional high-bypass turbofan engines. Since it does not have LOX on-board, it can meet all airport noise and safety standards, allowing it to operate from virtually any airport.
  • the RLV can either cruise for thousands of miles using JP-8 stored in its wing tanks, or prepare for launch. Once a decision to go for launch is made, LOX generation begins, and the fanjets are switched from JP-8 to gaseous hydrogen fuel. It should be noted that the gaseous hydrogen is .
  • the LOX generator consumes liquid hydrogen at rates between five and fifteen pounds per second, outputs approximately thirteen pounds of 90% LOX (the remaining 10% is inert gases) for every pound of LH 2 input, and then returns the input LH 2 at high pressure, in gaseous form, to the turbofan engines .
  • the LOX generation system generates all of the RLV's liquid oxygen at altitude.
  • LOX tanking can take as little as one hour if it starts at high engine-bleed rates at low altitude (e.g., 5,450 meters/18,000 feet), or as long as three hours if the process starts at low engine-bleed rates at high altitude (e.g., 9,090 meters/30,000, feet.) .
  • This decision can be made by the pilot, or by the mission director.
  • the system By the time the system reaches Mach 2 it is already above 30,300 meters (100,000 feet) and the dynamic pressure is below 100 pounds per square foot. At approximately Mach 8, the propellant cross-feeds disconnect, the first stage throttles back to match the acceleration .of the second stage, and the vehicles separate. The first stage then shuts down its engines and, using RCS, rotates to high angle of attack (- 65 degrees) foe reentry. The second stage proceeds to LEO and begins payload operations as required. The second stage is not impacted by the LOX production mechanism and is very similar to parallel- burn RLV second stages proposed before.
  • the first stage has such a large platform area, and relatively thick skins, it is able to reenter the atmosphere using the skins as a heat sink (i.e. without any additional Thermal Protection Systems (TPS) ) .
  • the preferred embodiment includes a TPS.
  • TPS Thermal Protection Systems
  • the ballistic coefficient is so low for reentry that first stage slows down to subsonic velocities well above 30,300 meters (100,000 feet).. Eliminating TPS on the first stage can be done to reduce to operations costs .
  • the first stage After reentry, the first stage restarts the high-bypass turbofans on JP-8 and flies to the nearest airport for a checkout and full, refueling for the trip back to base.
  • the description of the present invention has been made with respect to specific arrangements and constructions of a space transportation architecture. It will.be apparent to those skilled in the art that the foregoing description is for illustrative purposes only, and that various changes and modifications can be made to the present invention without departing from the overall spirit and scope of the present invention. The full extent of the present invention, is defined and limited only by the following claims.

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  • Engineering & Computer Science (AREA)
  • Remote Sensing (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Physics & Mathematics (AREA)
  • Plasma & Fusion (AREA)
  • Filling Or Discharging Of Gas Storage Vessels (AREA)

Abstract

A space transportation architecture includes an aerospaceplane (10) capable of horizontal take-off and landing, the aircraft including a means for generating liquid oxygen from ambient air while in flight and either an orbiter (18) or an expendable upper stage (20) coupled to the aerospaceplane.

Description

COMMERCIAL SPACE TRANSPORTATION SYSTEM
BACKGROUND OF THE INVENTION
1. Field of the Invention
The present invention relates to the field of aerospace vehicles, and more particularly, to a novel method and apparatus of placing a vehicle in low-earth orbit.
2. Prior Art .
It is known in the prior art that rockets can be used to launch space vehicles and to place satellites into earth orbit .
One of the most commonly used form of rockets is the "liquid fueled" rocket. In a liquid fueled rocket, a propellant mixes with an oxidizer and burns to provide thrust. All fuel must combine with an oxidizer in order to burn. At sea level, a separate oxidizer is not needed, as ambient air can serve as the oxidizer. At high altitudes (and in space) , however, the air thins and there is no readily available oxidizer which allows the fuel to burn. It is therefore necessary for the spacecraft to carry its own oxidizer.
The most common form of fuel is liquid hydrogen (referred to as LH2) , and the most common oxidizer is liquid oxygen (LOX) . Both LH2 and LOX must be cooled to very low temperatures in order to maintain their liquid state.
One of the drawbacks of the liquid fueled rocket is that a large volume of oxidizer must be carried to permit burning of the fuel. In a H2-LOX rocket, the ratio of oxidizer to fuel is typically 6 to 1. Thus, it is necessary for the rocket to carry a large supply of oxidizer for a given payload size. If this weight of oxidizer could be reduced before the spacecraft were to take off, it would greatly increase the overall efficiency of the spacecraft operation.
It is known in the prior art that oxidizer can be manufactured from ambient air as a spacecraft is flying. This prior art method, however, can only be used at high supersonic speeds. Flying a craft at supersonic speeds is generally not desirable.
The present invention overcomes the limitations of the prior art by providing a method and apparatus which permits a space vehicle to take off without an initial load of oxidizer. After take-off, the spacecraft cruises at a predetermined altitude at subsonic speed. The spacecraft carries an onboard liquid oxygen (LOX) generator. The LOX generator uses liquid hydrogen fuel to cool incoming air, and liquefy the gaseous oxygen. This liquid oxygen is then stored in a tank until it is needed to serve as an oxidizer in a rocket engine.
SUMMARY OF THE PRESENT INVENTION The present invention provides a novel space transportation architecture which overcomes the limitations of the prior art . The architecture includes an areospaceplane capable of horizontal take-off and landing. The areospaceplane includes a device for generating liquid oxygen from ambient air while in flight . Either an orbiter or an expendable upper stage can be coupled to the areospaceplane.
BRIEF DESCRIPTION OF THE DRAWINGS
Figure 1 is a perspective view showing an aerospaceplane of the present invention. Figure 2 is a perspective view showing an orbiter 12 of the present invention.
Figure 3 is a perspective view showing a cargo upper stage 14 of the present invention.
Figure 4 is a diagram of a typical mission scenario of a spacecraft employing the method of the present invention.
Figure 5 is a table which lists the major engine specifications used in the present invention.
Figure 6 illustrates an idealized view of the booster 16 used in the system of the present invention. Figure 7 is a plan view of the booster 16.
Figure 8 illustrates an idealized view of a second stage orbiter 18 used in the present invention.
Figure 9 is a plan view and an elevation of the orbiter 18. Figure 10 illustrates an expendable upper stage 20 which can be used in place of the orbiter.
Figure 11 shows the structural design of the upper stage 20.
DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT
An novel space transportation architecture will be described. In the following description, for the purposes of explanation, specific method steps, component arrangements and constructions and other details are set forth in order to provide a more thorough understanding of the present invention. It will be apparent to those skilled in the art, however, that the present invention may be practiced without these specific details. In other instances, well known manufacturing methods and structures have not been described in detail so as not to obscure the present invention unnecessarily.
The present invention allows a reusable space vehicle to attain low-earth orbit in a manner which requires the use of a minimum amount of fuel for a given amount of payload. The preferred embodiment of the present invention includes an areospaceplane as is shown in Figure 1. The areospaceplane is able to take off and land in a manner like a conventional airplane . The spacecraft therefore includes wings and horizontal stabilizers and a vertical stabilizer. The areospaceplane is a piloted vehicle, equipped with air breathing engines. The areospaceplane also includes a liquid oxygen (LOX) generation system. One such LOX generation system is described in United States Patent Application Serial No. 09/515,951, which is the preferred embodiment of the invention. It will be understood that other methods of generating LOX can be used with equal ef ectiveness. It is anticipated that the present invention can utilize wide variety of different mechanical means to compress and cool the air to form LOX.
The various components of the system of the present invention will now be described. Figure 6 shows an idealized form of the booster. In the preferred embodiment of the. present invention, the booster is capable of a total of Δv of
3.4 km/s (11,050 fps) , using five RL200 LOX/LH rocket engines.
The major engine specifications are listed in Figure 5. Rocket engine performance is expected to degrade by about 10% from the nominal specifications, due to the adjustment in oxidizer chemistry required by the LOX generation and propulsion system.
The booster's airframe in the preferred embodiment is a combination of aluminum and composite structural elements .
The total wing in the preferred embodiment is 511 m2 (5500 sq. ft) , but it will be apparent to those of skill in the art that a larger or smaller wing planform can be used. Due to the LOX tank being empty at take-off as well as landing, the vehicle is capable of utilizing a standard Boeing 777 landing gear.
Figure 7 shows a plan view of the preferred design for the booster.
In addition to its five rocket engines, the vehicle is equipped with four turbo-fan engines, modified to operate on standard jet fuel as well as the gaseous hydrogen fuel provided by the LOX generation system. The system's second stage orbiter is shown in Figure 8. The second stage is a pressurized orbital vehicle, capable of transporting cargo as well as human crew. The preferred design uses two rocket - engines (which, in the preferred embodiment are identical to those on the booster) for its main propulsion system. This gives the orbiter a total anticipated Δv capability of 5.5 km/s (18,150 fps) . In addition, it is equipped with two Orbital Maneuvering System (OMS) engines and Attitude Control System (ACS) thrusters, all of which will operate on non-toxic (green) storable propellants.
The airfra e of the orbiter is constructed entirely of composites, including an integral forward crew compartment capable of transporting up to 8,165 kg (18,000 lbs.) to and from orbit. The wing surface area in the preferred embodiment is approximately 93 m2 (1000 sq. ft) . Figure 9 shows a top and side view of the design.
Figure 10 illustrates an expendable upper stage which can be used in place of the orbiter of Figure 8. The upper stage is an unpressurized orbital vehicle intended for transporting large and heavy payloads, such as GEO satellites and ISS-class habitation modules. It can be used in a reusable configuration utilizing a deployable recovery system, or as a completely. second stage in order to increase the delivered payload capability. The upper state uses two rocket engines similar to those in the booster, together with a non- oxic RCS. It does not include a separate OMS in its current design. The structural design of the Upper Stage is shown in Figure 11. It incorporates a 18.3 (60 ft) payload fairing, capable of supporting up to 27,215 kg (60,000 lbs.).
A typical mission profile of the commercial space transportation architecture of the present invention will now be described. In this example the orbiter is first mated to the areospaceplane. The areospaceplane then takes off from a traditional airport facility, and climbs to a cruising altitude of approximately 25K feet. The LOX system is then used to generate fuel for the rocket engine. After a sufficient amount of LOX has been generated, the rocket engine is used to lift the areospaceplane to altitude. At a speed of approximately Mach 8, the vehicles separate. The orbiter continues to climb until it reaches the desired low earth orbit . The areospaceplane reenters the atmosphere and lands at the nearest airport .
A more detailed description of a baseline mission scenario using the system of the present invention will now be given. A schematic diagram of the mission profile is shown in Figure 12. The system takes off with a compliment of LH2 and JP-8 and climbs using conventional high-bypass turbofan engines. Since it does not have LOX on-board, it can meet all airport noise and safety standards, allowing it to operate from virtually any airport. At altitude, the RLV can either cruise for thousands of miles using JP-8 stored in its wing tanks, or prepare for launch. Once a decision to go for launch is made, LOX generation begins, and the fanjets are switched from JP-8 to gaseous hydrogen fuel. It should be noted that the gaseous hydrogen is . a byproduct from the preferred embodiment of the system used to generate LOX, which will be described in the next section. The LOX generator consumes liquid hydrogen at rates between five and fifteen pounds per second, outputs approximately thirteen pounds of 90% LOX (the remaining 10% is inert gases) for every pound of LH2 input, and then returns the input LH2 at high pressure, in gaseous form, to the turbofan engines .
In this manner, the LOX generation system generates all of the RLV's liquid oxygen at altitude. LOX tanking can take as little as one hour if it starts at high engine-bleed rates at low altitude (e.g., 5,450 meters/18,000 feet), or as long as three hours if the process starts at low engine-bleed rates at high altitude (e.g., 9,090 meters/30,000, feet.) . This decision can be made by the pilot, or by the mission director. Once LOX tanking is finished, the system assumes the proper heading, all rocket engines fire, and the combined vehicles begin a rapid climb. The turbofan engines are shutdown about Mach 1.4 and the inlets covered. By the time the system reaches Mach 2 it is already above 30,300 meters (100,000 feet) and the dynamic pressure is below 100 pounds per square foot. At approximately Mach 8, the propellant cross-feeds disconnect, the first stage throttles back to match the acceleration .of the second stage, and the vehicles separate. The first stage then shuts down its engines and, using RCS, rotates to high angle of attack (- 65 degrees) foe reentry. The second stage proceeds to LEO and begins payload operations as required. The second stage is not impacted by the LOX production mechanism and is very similar to parallel- burn RLV second stages proposed before.
Because the first stage has such a large platform area, and relatively thick skins, it is able to reenter the atmosphere using the skins as a heat sink (i.e. without any additional Thermal Protection Systems (TPS) ) . The preferred embodiment, however, includes a TPS. In particular, there is 0.2 kg per square meter (1.0 pounds per square foot) of TPS on the lower 60% of the first stage. The ballistic coefficient is so low for reentry that first stage slows down to subsonic velocities well above 30,300 meters (100,000 feet).. Eliminating TPS on the first stage can be done to reduce to operations costs . After reentry, the first stage restarts the high-bypass turbofans on JP-8 and flies to the nearest airport for a checkout and full, refueling for the trip back to base. The description of the present invention has been made with respect to specific arrangements and constructions of a space transportation architecture. It will.be apparent to those skilled in the art that the foregoing description is for illustrative purposes only, and that various changes and modifications can be made to the present invention without departing from the overall spirit and scope of the present invention. The full extent of the present invention, is defined and limited only by the following claims.

Claims

CLAIMSWhat is claimed is:
1. A space transportation architecture, comprising: an aerospaceplane capable of horizontal take-off and landing, said aerospaceplane including a means for generating liquid oxygen from ambient air while in flight; a spacecraft coupled to said aerospaceplane.
2. A space transportation architecture, comprising: an aerospaceplane capable of horizontal take-off and landing, said aerospaceplane including a means for generating liquid oxygen from ambient air while in flight; an expendable upper stage coupled to said aerospaceplane.
3. A method of placing a vehicle in orbit comprising: providing an architecture as claimed in claim 1 or claim 2, causing the aerospaceplane to take off, generating liquid oxygen while in flight, and launching the vehicle into orbit from the aerospaceplane.
PCT/GB2003/001582 2002-04-10 2003-04-10 Commercial space transportation system WO2003086860A1 (en)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US12007802A 2002-04-10 2002-04-10
US10/120,078 2002-04-10

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WO2003086860A9 true WO2003086860A9 (en) 2004-02-12

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US6283227B1 (en) 1998-10-27 2001-09-04 Schlumberger Technology Corporation Downhole activation system that assigns and retrieves identifiers
CA2870808C (en) * 2012-04-04 2021-01-26 Commercial Aerospace Plane Pty Limited An aerospace plane system
US10384796B2 (en) 2012-04-04 2019-08-20 Commercial Aerospace Plane Pty Limited Aerospace plane system
RU2529121C2 (en) * 2012-10-16 2014-09-27 Виктор Степанович Ермоленко Two-stage aerospace system (versions)

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US3756024A (en) * 1962-02-23 1973-09-04 Gen Dynamics Corp Method and apparatus for coordinating propulsion in a single stage space flight
US6119985A (en) * 1997-03-07 2000-09-19 Pioneer Rocketplane Corporation Reusable rocket-propelled high altitude airplane and method and apparatus for mid-air oxidizer transfer to said airplane
FR2766911B1 (en) * 1997-08-04 1999-09-03 Air Liquide METHOD AND APPARATUS FOR TRANSFERRING MATERIAL BETWEEN A LIQUID AND A GAS FLOWING COUNTER-CURRENT, AND VEHICLE EQUIPPED WITH SUCH APPARATUS
US6193187B1 (en) * 1998-12-31 2001-02-27 Harry Scott Payload carry and launch system
JP2001317406A (en) * 2000-02-29 2001-11-16 Andrews Space & Technology Inc Method and device for placing artificial satellite into low earth orbit

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