Abstract:
A reusable unmanned single-stage boost-glide suborbital Earth launcher, able to propel large or smaller payloads from Earth surface or from air launch, to a nearly Earth-orbital condition, then glide circum-globally to horizontal airstrip landing at its launch site or similar circum-global recovery site. The payload, using its own propulsion, is thus enabled to complete insertion into low-Earth-orbit (LEO) or destinations beyond LEO. This technically feasible capability, beyond conferring economic benefits of reusability and maneuverability, can be adapted to increase payload by adding drop tanks, or solid or liquid “strap on” boosters, and might eventually be modified and evolved to perform manned aerospace missions, as well as single stage to orbit (SSTO) missions.

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
       [0001]    This application claims priority to and the benefit of the filing of U.S. Provisional Patent Application Ser. No. 61/793,441, entitled “Reusable Global Launcher”, filed on Mar. 15, 2013, the specification and claims of which are incorporated herein by reference. 
     
    
     BACKGROUND OF THE INVENTION 
       [0002]    1. Field of the Invention (Technical Field) 
         [0003]    Embodiments of the present invention are related to Earth or air launch reusable rocket vehicle systems. 
         [0004]    2. Background of the Invention 
         [0005]    Note that the following discussion refers to a number of publications and references. Discussion of such publications herein is given for more complete background of the scientific principles and is not to be construed as an admission that such publications are prior art for patentability determination purposes. 
         [0006]    Previous designs for reusable rocket systems include: 1) Earth liftoff boost stages which execute short-range turnaround return to the launch site or a downrange site by ballistic, glide or powered flight with parachutes, vertical-touchdown or horizontal airstrip recovery; 2) Earth liftoff single-stage-to-orbit (SSTO) launchers to low-Earth-orbit (LEO) which return to any Earth site by ballistic, glide or powered flight with parachute, vertical-touchdown or horizontal airstrip recovery: 3) air-launch for any of the above. The SSTO option has long been dismissed as “too technologically risky” because of vehicle development dry weight uncertainty. 
       SUMMARY OF THE INVENTION 
       [0007]    An embodiment of the present invention is a method for operating a launch vehicle, the method comprising the steps of launching a reusable launch vehicle; releasing an external vehicle payload from the vehicle at a velocity lower than orbital velocity; the vehicle subsequently performing global flight; and landing the vehicle. The vehicle is preferably unmanned. The landing step is optionally performed at the same location as the launch step. The launching step optionally comprises air launching. The vehicle preferably comprises sufficient aerodynamic glide capability to circle the Earth, preferably having an L/D ratio of greater than or equal to approximately 3. The vehicle optionally performs an aerodynamic dog-leg maneuver during ascent prior to the releasing step. The vehicle preferably comprises a rocket. The landing step preferably comprises a landing horizontally on an airstrip. The vehicle is preferably a single stage vehicle. The vehicle preferably comprises one or elements selected from the group consisting of cylindrical propellant tanks, expendable drop tanks which preferably supply fuel to the vehicle to increase its propulsion and/or thrust, no internal cargo bay, duel fuel propulsion, circular cross-section, dry wings, and extra engines. The method optionally further comprises the payload entering low earth orbit and/or destinations beyond low earth orbit. 
         [0008]    Objects, advantages and novel features, and further scope of applicability of the present invention will be set forth in part in the detailed description to follow, taken in conjunction with the accompanying drawings, and in part will become apparent to those skilled in the art upon examination of the following, or may be learned by practice of the invention. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0009]    The accompanying drawings, which are incorporated into and form a part of the specification, illustrate one or more embodiments of the present invention and, together with the description, serve to explain the principles of the invention. The drawings and the dimensions therein are only for the purpose of illustrating certain embodiments of the invention and are not to be construed as limiting the invention. In the drawings: 
           [0010]      FIG. 1  shows a cycle schematic of a DF/DX example engine. 
           [0011]      FIG. 2   a  is a side view of an embodiment of the reusable single-stage global launcher of the present invention in the horizontal landing configuration. 
           [0012]      FIG. 2   b  is a top view of the embodiment of the reusable single-stage global launcher of the present invention shown in  FIG. 2   a  with a payload container. 
           [0013]      FIG. 2   c  is a side view of the embodiment of the reusable single-stage global launcher of the present invention shown in  FIG. 2   b.    
           [0014]      FIG. 3   a  is a top view of another embodiment of the reusable single-stage global launcher of the present invention with a single payload container. 
           [0015]      FIG. 3   b  is a top view of another embodiment of the reusable single-stage global launcher of the present invention with a single payload container and two external fuel tanks. 
           [0016]      FIG. 3   c  is a side view of the embodiment of the reusable single-stage global launcher of the present invention shown in  FIG. 3   b.    
           [0017]      FIG. 4   a  is a top view of another embodiment of the reusable single-stage global launcher of the present invention with a single payload container. 
           [0018]      FIG. 4   b  is a top view of another embodiment of the reusable single-stage global launcher of the present invention with a single payload container and two external fuel tanks. 
           [0019]      FIG. 4   c  is a side view of the embodiment of the reusable single-stage global launcher of the present invention shown in  FIG. 4   b.    
           [0020]      FIG. 5  compares the sizes of various vehicles.  FIG. 5   a  shows a Saturn V rocket.  FIG. 5   b  shows an Energia SL-17.  FIG. 5   c  shows the reusable global launcher of  FIG. 2 .  FIG. 5   d  shows the reusable global launcher of  FIG. 3 .  FIG. 5   e  shows AN-225 aircraft. 
           [0021]      FIG. 6  compares the sizes of various vehicles.  FIG. 6   a  shows a Delta IV Heavy launcher.  FIG. 6   b  shows a 1½-stage vehicle with a 16.5 ft. main tank diameter and 2xRD-180 plus 2xSSME engines.  FIG. 6   c  shows a 1½-stage vehicle with a 16.5 ft main tank diameter and 5xDF/DX engines.  FIG. 6   d  shows a 1½-stage vehicle with a 12.5 ft. main tank diameter and 1 xRD-180 plus 1xSSME engines.  FIG. 6   e  shows a 747-400 aircraft. 
           [0022]      FIG. 7   a  graphs comparisons of vehicle dry mass estimates of embodiments of the present invention vs. vehicle total estimated volume. 
           [0023]      FIG. 7   b  shows a proposed reference dry mass equation relating vehicle dry mass to total vehicle volume. 
           [0024]      FIG. 8  shows example reusable global launcher (RGL) performance analysis, with and without an upper “kick” stage, for cargo delivery to LEO. 
           [0025]      FIG. 9   a  is a side view of an embodiment of the reusable global launcher of the present invention as the Earth launch platform for a geolunar shuttle. 
           [0026]      FIG. 9   b  is a top view of the geolunar shuttle shown in  FIGS. 9   a  and  9   c.    
           [0027]      FIG. 9   c  is a top view of an embodiment of the reusable global launcher/geolunar shuttle combination shown in  FIG. 9   a.    
           [0028]      FIG. 10   a  depicts top, rear, and side views of an embodiment of the reusable global launcher of the present invention suspended from the underside of a specially designed subsonic launch aircraft. 
           [0029]      FIG. 10   b  shows a reference AN-225 aircraft for size comparison with the aircraft of  FIG. 10   a.    
           [0030]      FIG. 10   c  is a detail of the embodiment of the reusable global launcher shown in  FIG. 10   a.    
           [0031]      FIG. 11   a  shows top and side views of another embodiment of a reusable global launcher of the present invention suspended from the underside of another, scaled up subsonic launch aircraft. 
           [0032]      FIG. 11   b  is a rear view of the vehicle shown in  FIG. 11   a.    
           [0033]      FIG. 11   c  shows a reference AN-225 aircraft for size comparison with the aircraft of  FIG. 11   a.    
       
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
       [0034]    Embodiments of the present invention introduce the novel possibility of an unmanned reusable single-stage or 1½ stage global launcher, which may either be be large or small enough to be air-launched, which can circle the Earth without attaining LEO, and glide to its launch or other Earth sites, by incorporating in its design, a high (approximately 3) hypersonic lift: drag (L/D) configuration. Such a configuration enables the launcher to return to its launch site or other sites from a suborbital or unsustainable Earth orbital (UEO) condition. As used throughout the specification and claims, the term “launch vehicle” or “launcher” means a vehicle that comprises a thrust to weight ratio to sufficient to lift off from the surface of the earth. As used throughout the specification and claims, the terms “unsustainable Earth orbit” or “UEO” mean a temporary Earth orbit with a perigee low enough that it soon decays. A 50×100 nautical mile orbit is often used as a reference UEO and will be used as an example herein. As used throughout the specification and claims, the term “global flight” means unpowered flight at least once around the Earth without starting from or attaining low Earth orbit, including but not limited to unpowered flight in a UEO. Global flight may optionally include gliding flight segments and/or ballistic flight segments, such as those occurring during skip-gliding. As used throughout the specification and claims, the term “skip-gliding” means flight that alternates between ballistic trajectory phases and phases in which aerodynamic forces are large enough to provide lift. Skip gliding has never been performed by a launch vehicle. 
         [0035]    Embodiments of the present invention comprise (i) vehicle means and techniques for improved launch from Earth surface or airborne platforms, for a range of large or small manned or unmanned payloads, and including the economic benefits of launch vehicle recovery and reusability; (ii) a reusable Earth launch vehicle system having sufficient aerodynamic glide capability to complete at least one Earth circuit from a suborbital condition, to ensure recovery at its launch site or geographically similar site; (iii) a reusable Earth launch vehicle system having sufficient aerodynamic maneuvering capability to achieve large offsets during ascent and return, to increase flexibility and to complicate interception; and (iv) a reusable Earth launch vehicle system having sufficient aerodynamic glide and maneuvering capability to ensure at least “once-around” return and recovery, while accommodating expanded ranges of uncertainty in launch vehicle dry weight during development. 
         [0036]    The above capability preferably confers the following new advantages: 1) by virtue of high hypersonic L/D configuration, the launcher can “dog-leg” its heading during ascent to offset location and azimuth of payload release as well as during return, to increase flexibility and to confuse attempted interception; 2) by virtue of its high L/D configuration, and consequent extended global glide range to ensure return to global recovery, overall aerodynamic performance can provide wider margin allowances for dry weight growth during vehicle development. 
         [0037]    Embodiments of the present invention are directed to a launch system including process and apparatus for launch from Earth surface or airborne platforms, return from at least one Earth circuit recovery and reuse of a launch vehicle. A single-stage rocket propulsive vehicle is contemplated, which launches vertically or from an airborne platform, carries its payload to a suborbital or UEO release condition in any location, then can return by unpowered glide to complete at least one Earth circuit to its launch site or similar geographic location, thereby eliminating the typical requirement of having to turn around and fly back to the original launch site. 
         [0038]    For simplicity leading to development, performance, operational and overall economic advantage, the reusable launcher is preferably unmanned, carries all main propellants in cylindrical tankage, and is capable of unpowered glide to horizontal airstrip landing. Accordingly, the launcher is preferably configured as “dart-shaped,” having a cylindrical body and highly swept platform, with a sweep angle of approximately 81 degrees. Configuration and aerodynamic feature of this high-fineness vehicle are disclosed in U.S. Pat. No. 5,090,642. Such a vehicle having hypersonic L/D of approximately 3.0, injected horizontally at a re-entry velocity of 26,000 feet/second and altitude of 170,000 feet, could glide more than 11,500 nautical miles directly downrange and ±3,500 nautical miles crossrange, with horizontal airstrip recovery. Use of skip-gliding may extend and maximize these ranges. 
         [0039]    Cylindrical tankage technology for this vehicle is established from: 1) Saturn V (33 feet); 2) NASA shuttle, (27.5 feet); 3) Delta IV (16.5 feet); Atlas V (12.5 feet). Propulsion technology for this vehicle is established from: 1) NASA space shuttle main engine (SSME/RS-25D); 2) Atlas V (RD-180); and projected from Russia (RD-701) and US (DF/DX). Typical specification for these developed and projected engines are presented in Table 1, summarized from “Dual-Fuel Propulsion in Single-Stage Advanced Manned Launch System Vehicle”, Lepsch, R. A. Jr, et al, NASA Langley Research Center, AIAA Journal of Spacecraft and Rockets, Vol. 32, No. 3, May-June 1995. Though RD-180 and SSME with decades of successful flight experience may be considered obsolete, they do represent what can be achieved with demonstrated and foreseeable rocket engine technology. 
         [0000]    
       
         
               
             
               
               
               
             
               
               
               
               
             
               
               
               
               
               
               
             
               
               
               
               
               
               
             
           
               
                 TABLE 1 
               
             
             
               
                   
               
               
                 Rocket Engine Data 
               
             
          
           
               
                   
                 SINGLE-FUEL (EXISTING) 
                   
               
             
          
           
               
                   
                   
                 SPACE SHUTTLE MAIN 
                 DUAL FUEL 
               
               
                   
                   
                 ENGINE (SSME-O 2 /H 2 ) 
                 (DUAL EXPANDER DESIGN) 
               
             
          
           
               
                 ENGINE 
                 RD-180 
                 STD.  
                 MOD.  
                 MODE I 
                 MODE 2 
               
               
                 PARAMETER 
                 (O 2 /KEROSENE) 
                 (ε  =  77.5) 
                 (ε = 50)  
                 (O 2 /PROPANE/H 2 ) 
                 (O 2 /H 2 ) 
               
               
                   
               
             
          
           
               
                 Thrust, sea level, lbf 
                 859,800 
                 391,900 
                 420,000 
                 666,700 
                 N/A 
               
               
                 Thrust, vacuum, lbf 
                 934,800 
                 488,300 
                 482,200 
                 750,000 
                 235,100 
               
               
                 Specific impulse, sl., sec  
                 311.3 
                 363.3 
                 389.2 
                 341 
                 N/A 
               
               
                 Specific impulse, vac., sec 
                 337.8 
                 452.7 
                 447.0 
                 383.7 
                 462.9 
               
               
                 Chamber pressure, psia 
                 3,870 
                 3,126 
                 3,126 
                 5,000/2,500 
                 2,500 
               
               
                 Oxidizer: Fuel ratio 
                 2.6 
                 6.0 
                 6.0 
                 3.2/6.0 
                 6.0 
               
               
                 Nozzle expansion ratio, ε 
                 75.0 
                 77.5 
                 50.0 
                 74.8/36.3 
                 119.9 
               
               
                 Nozzle exit diem., ft 
                 4.91 
                 7.88 
                 6.33 
                 6.36 
                 6.36 
               
               
                 Engine length, ft 
                 11.8 
                 14.67 
                 13.00 
                 10.7 
                 10.7 
               
               
                 Engine dry mass, Ibm 
                 11,905 
                 7,203 
                 6,780 
                 8,127 
                 8,127 
               
               
                   
               
             
          
         
       
     
         [0040]    Payload capacity of this single stage launcher can be increased by addition of expendable drop tanks and extra engines on the reusable core vehicle to provide required start thrust. Augmented with drop tanks, this version of the launcher will be termed “1½-stage”. 
         [0041]    For heavy versions of this vehicle, i.e. main tank diameter between approximately 27.5 feet and approximately 33.0 feet or greater, reduced payloads could be carried farther directly into LEO. Doing so, however, would impose requirements for additional on-board propulsion systems for orbital maneuvering and de-orbit for Earth return. Rather, jettisoning the container at suborbital altitude and velocity, and continuing from there with a separate upper “kick” stage would allow automatic Earth return of the launcher and deliver significantly increased cargo to LEO or far space transfer, because of the classical benefits of staging. 
         [0042]    Embodiments of the invention preferably incorporate and consolidate performance, economic and operational advantages of five individual system concepts: 1) economic benefit of reusability; 2) simplicity of unmanned operation: 3) simplicity of circular cross-section tube-and-dry-wing design with no internal cargo bay; 4) aerothermodynamic benefit of high-fineness configuration; and 5) launch performance benefit of dual-fuel propulsion. 
         [0043]    Embodiments of the present invention can incorporate single-fuel (oxygen-hydrogen) as well as dual-fuel (oxygen-hydrocarbon-hydrogen) propulsion. Comparison shows that in all cases, dual-fuel propulsion outperforms oxygen-hydrogen propulsion in terms of payload: gross weight and payload: dry weight ratios. The dual-fuel/dual-expander (DF/DX) engine example used herein is from Lepsch et al, adapted from R. Beichel, “Dual Fuel Dual Expander Rocket Engine”, U.S. Pat. No. 4,220,001, descended from R. Salkeld, “Mixed-Mode Propulsion Aerospace Vehicles”, U.S. Pat. No. 3,955,784. That DF/DX example engine, a cycle schematic of which is shown in  FIG. 1 , combines hydrogen and hydrocarbon combustion within the same engine using coannular combustion chambers, which exhaust through a common nozzle. The outer chamber burns oxygen and hydrogen while the inner chamber burns oxygen and hydrocarbon fuel. As with the RD-701, the engine operates initially in a dual-fuel mode, where both hydrocarbon and hydrogen are burned, and then transitions to a single-fuel mode where only hydrogen is burned. During the single-fuel mode, the nozzle expansion ratio is effectively increased due to the reduction in throat area when the hydrocarbon chamber is inoperative. This results in high performance at altitude. A number of hydrocarbon propellant types have been investigated for use in this concept; subcooled liquid propane was selected for this study in view of results showing that lower vehicle empty weight can be achieved with this hydrocarbon fuel than with other common ones. A gas generator cycle is assumed for the hydrocarbon portion of the engine, and a staged combustion cycle is assumed for the hydrogen portion. Dual-fuel volume splits within the following core vehicles and drop tanks are thought to be near optimum; formal optimization would increase vehicle performance benefits of the dual-fuel vs. single-fuel options. 
         [0044]      FIG. 2  depicts an embodiment of the reusable single-stage global launcher of the present invention with a main tank diameter of 33.0 feet using Saturn 5 main tankage technology heritage and flight-proven RD-180 and SSME propulsion technology. Parameters for this embodiment are listed in Table 2. 
         [0000]    
       
         
               
             
               
               
               
             
           
               
                 TABLE 2 
               
               
                   
               
               
                 Vehicle Parameters 
               
               
                   
               
             
             
               
                   
               
             
          
           
               
                   
                 Payload a   
                 302,000 
               
               
                   
                 Cargo, lbm 
                 273,500 
               
               
                   
                 Cargo container, lbm 
                 28,500 
               
               
                   
                 Gross liftoff mass, lbm 
                 7,607,200 
               
               
                   
                 Dry mass less engines, lbm b   
                 368,500 
               
               
                   
                 Engines, lbm 
                   
               
               
                   
                 8 × RD-180 
                 95,240 
               
               
                   
                 6 × SSME(ε = 77.5) 
                 43,218 
               
               
                   
                 Re-entry planform loading, lbm/ft 2   
                 228.2 
               
               
                   
                 Cargo space, ft 
                 22 × 158 
               
               
                   
                 Cargo density, lbm/ft 3   
                 4.8 
               
               
                   
                 Return glide downrange, n.mi. 
                 (global) 
               
               
                   
                 Return glide crossrange, n.mi. 
                 ±3,500 
               
               
                   
                   
               
               
                   
                   a 50 × 100 n.mi, 28.7°,  b incl. 15% margin 
               
             
          
         
       
     
         [0045]      FIG. 3  depicts an embodiment of the reusable global launcher of the present invention with a main tank diameter of 27.5 feet using NASA shuttle external tankage technology heritage and flight-proven RD-180 and SSME propulsion technology. At this tank diameter, the designated high-pressure propulsion can provide the required single-stage liftoff thrust without occupying all of the available vehicle base area. Therefore, additional engines and external propellant drop tanks can be added to result in a heavier 1½-stage design, with payload/cargo increase of about 100 percent. Vehicle parameters for this embodiment are listed in Table 3. 
         [0000]    
       
         
               
               
               
             
               
               
               
               
             
               
               
               
               
             
           
               
                 TABLE 3 
               
             
             
               
                   
               
               
                   
                   
                 1 ½ - STAGE 
               
             
          
           
               
                 VEHICLE 
                 SINGLE 
                   
                 DROP  
               
               
                 PARAMETER 
                 STAGE 
                 CORE 
                 TANKS (2) 
               
               
                   
               
             
          
           
               
                 Payload a   
                 158,000 
                 318,000 
                 — 
               
               
                 Cargo, lbm 
                 143,100 
                 294,200 
                 — 
               
               
                 Cargo container, lbm 
                 14,900 
                 23,800 
                 — 
               
               
                 Gross liftoff mass, lbm 
                 4,515,000 
                 4,726,000 
                 1,561,400 
               
               
                 Dry mass less engines, lbm b   
                 242,190 
                 260,900 
                 62,900 
               
               
                 Engines, lbm 
                   
                   
                   
               
               
                 5 (SS) or 7 (1.5 S) × RD = 180 
                 59,525 
                 83,335 
                 — 
               
               
                 3 (SS) or 4 (1.5 S) × SSME 
                 21,609 
                 28,812 
                 — 
               
               
                 (ε = 77.5) 
                   
                   
                   
               
               
                 Re-entry planform loading,  
                 26.3 
                 30.2 
                 — 
               
               
                 lbm/ft 2   
                   
                   
                   
               
               
                 Cargo space, ft 
                 17 × 122 
                 21 × 155 
                 — 
               
               
                 Cargo density, lbm/ft 3   
                 5.5 
                 5.8 
                 — 
               
               
                 Return glide downrange, n.mi. 
                 (global) 
                 (global) 
                 — 
               
               
                 Return glide crossrange, n.mi. 
                 ±3,500 
                 ±3,500 
                 — 
               
               
                   
               
               
                   a 50 × 100 n.mi.; 28.7°,  b incl. 15% margin 
               
             
          
         
       
     
         [0046]      FIG. 4  shows an embodiment of the reusable global launcher of the present invention having a main tank diameter of 33.0 feet using Saturn 5 main tankage technology heritage and projected DF/DX propulsion technology. Growth to 1½-stage design are as described for  FIGS. 2 and 3 . Vehicle parameters for this embodiment are listed in Table 4. 
         [0000]    
       
         
               
               
               
             
               
               
               
               
             
               
               
               
               
             
           
               
                 TABLE 4 
               
             
             
               
                   
               
               
                   
                   
                 1 ½ - STAGE 
               
             
          
           
               
                 VEHICLE 
                 SINGLE 
                   
                 DROP  
               
               
                 PARAMETER 
                 STAGE 
                 CORE 
                 TANKS (2) 
               
               
                   
               
             
          
           
               
                 Payload a   
                 373,000 
                 735,000 
                 — 
               
               
                 Cargo, lbm 
                 344,500 
                 700,700 
                 — 
               
               
                 Cargo container, lbm 
                 28,500 
                 34,300 
                 — 
               
               
                 Gross liftoff mass, lbm 
                 7,410,100 
                 7,834,800 
                 2,731,300 
               
               
                 Dry mass less engines, lbm b   
                 361,800 
                 382,400 
                 109,900 
               
               
                 Engines, lbm 
                   
                   
                   
               
               
                 14 (SS) or 19 (1.5 S) × DF/DX 
                 113,778 
                 154,473 
                 — 
               
               
                 Re-entry planform loading, lbm/ft 2   
                   
                   
                   
               
               
                 Cargo space, ft 
                 21 × 135 
                 25 × 188 
                 — 
               
               
                 Cargo density, lbm/ft 3   
                 6.1 
                 7.9 
                 — 
               
               
                 Return glide downrange, n.mi.  
                 11,500 
                 11,500 
                 — 
               
               
                 Return glide crossrange, n.mi. 
                 ±3,500 
                 ±3,500 
                 — 
               
               
                   
               
               
                   a 50 × 100 n.mi.; 28.7°,  b incl. 15% margin 
               
             
          
         
       
     
         [0047]      FIG. 5  shows the reusable global launchers from  FIGS. 2 and 3  compared with reference large aerospace vehicles, including AN-225, currently the world&#39;s largest aircraft. Capacities are in Table 5. 
         [0000]    
       
         
               
               
               
               
             
               
               
               
               
               
               
             
           
               
                 TABLE 5 
               
             
             
               
                   
               
               
                   
                 REFERENCE HEAVY 
                 REUSABLE GLOBAL LAUNCHER 
                 REFERENCE 
               
               
                   
                 LAUNCHER 
                 CONCEPTS 
                 AIRCRAFT 
               
             
          
           
               
                   
                 SATURN 5 
                 ENERGIA SL-17 
                 33 FT. DIA. 
                 27.5 FT. DIA. 
                 AN-225 
               
               
                   
                 1966-1975 
                 1987-1988 
                 TANK SINGLE 
                 TANK 1 ½ STAGE 
                 1992- 
               
               
                   
                 (FIG. 5a) 
                 (FIG. 5b) 
                 STAGE (FIG. 5c) 
                 (FIG. 5d) 
                 (FIG. 5e) 
               
               
                   
               
               
                 Payload weight lbm 
                   250,000 a   
                   230,000 b   
                   302,000 c   
                   318,000 c   
                   550,000 e   
               
               
                 Gross liftoff wt. lbm  
                 6,200,000 
                 5,357,000 
                 7,607,000 
                 6,287,000 
                 1,250,000 
               
               
                 Dry wt. lbm 
                   435,900 
                   497,000 
                   507,000 
                   436,000 d   
                   600,000? 
               
               
                   
               
               
                   a 100 n.mi., 27.8°,  b &lt;50 × 100 n.mi., 51.6°;  c 50 × 100 n.mi., 27.8°,  d Incl. drop tanks;  e 1,500 n.mi. range 
               
             
          
         
       
     
         [0048]      FIG. 6  shows three smaller reusable global launchers compared with reference aerospace vehicles including Delta IV Heavy, currently the world&#39;s largest operational launcher ( FIG. 6   a ). All are 1½-stage, since at these size ranges, single-stage reusable concepts are not credible. These three are: 1) with 16.5 ft. main tank diameter and 2=RD-180 plus 2xSSME engines ( FIG. 6   b ); 2) with 16.5 ft main tank diameter and 5xDF/DX engines ( FIG. 6   c ); 3) with 12.5 ft. main tank diameter and 1xRD-180 plus 1xSSME engines ( FIG. 6   d ). Capacities are listed in Table 6. 
         [0000]    
       
         
               
               
               
               
             
               
               
               
               
               
               
             
           
               
                 TABLE 6 
               
             
             
               
                   
               
               
                   
                   
                 REUSABLE GLOBAL LAUNCHER  
                   
               
               
                   
                 REFERENCE  
                 CONCEPTS 
                 REFERENCE 
               
             
          
           
               
                   
                 LAUNCHER 
                 16.5 FT. DIA. 
                 16.5 FT. DIA. 
                 27.5 FT. DIA. 
                 AIRCRAFT 
               
               
                   
                 DELTA IV HEAVY  
                 TANK 1 ½  
                 TANK 1 ½  
                 TANK 1 ½  
                 747-400 ER 
               
               
                   
                 2004- 
                 STAGE 
                 STAGE 
                 STAGE 
                 1969- 
               
               
                   
                 (FIG. 6a) 
                 (FIG. 6c) 
                 (FIG. 6c) 
                 (FIG. 6d) 
                 (FIG. 6e) 
               
               
                   
               
               
                 Payload weight, lbm 
                   52,900 a   
                   98,000 b   
                   158,000 b   
                   45,000 b   
                  60,740 d   
               
               
                 Gross liftoff wt., lbm  
                 1,632,600 
                 2,133,000 
                 2,520,000 
                 1,043,000 
                 910,000 
               
               
                 Dry wt., lbm 
                   170,000 
                   156,900 c   
                   172,600 c   
                   86,400 c   
                 406,900 
               
               
                   
               
               
                   a 250 n.mi., 28.7°;  b 50 × 100 n.mi., 28.7°;  c incl. drop tanks;  d 6,660 n.mi. range 
               
             
          
         
       
     
         [0049]    Tables 7-9 list itemized mass and payload estimates for reusable global launchers with main tank diameters 33.0, 27.5, and 16.5 and 12.5 feet respectively, for payloads injected into 50×100 nautical mile UEO. Eighteen design options are presented including single-fuel oxygen/hydrogen, as well as dual-fuel oxygen/hydrocarbon/hydrogen propulsion. In all comparisons, dual-fuel outperforms single-fuel propulsion. Dry mass estimates assume aluminum-lithium orthogrid-stiffened cylindrical tankage and appropriate use of graphite epoxy composite materials in airframe structures. Use of composites in tankage would reduce dry mass and increase vehicle performance. 
         [0000]    
       
         
               
             
               
               
             
               
               
               
             
               
               
               
               
               
               
               
             
               
               
               
               
               
               
               
             
           
               
                 TABLE 7 
               
             
             
               
                   
               
               
                 Vehicle Mass Estimates (Main Tank Diam: 33 feet) 
               
             
          
           
               
                   
                 VEHICLE 
               
             
          
           
               
                   
                 FULLY REUSABLE 
                 GROWTH WITH O 2 /H 2  DROP 
               
               
                   
                 (Single Stage) 
                 TANKS (1½ - Stage) 
               
             
          
           
               
                   
                 8xRD-180 
                   
                   
                 10xRD-180 
                   
                   
               
               
                   
                 6xSSME 
                 15xSSME 
                   
                 6xSSME 
                 19xSSME 
               
               
                 WEIGHTS 
                             = 77.5 
                             = 50.0 
                 14xDF/DX 
                             = 77.5 
                             = 50.0 
                 19xDF/DX 
               
               
                   
               
             
          
           
               
                 Tankage 
                 102,600 
                 78,300 
                 99,100 
                 102,600 
                 94,700 
                 99,100 
               
               
                 Body incl. thrust structure 
                 62,700 
                 42,400 
                 63,700 
                 65,800 
                 57,600 
                 68,700 
               
               
                 Wing group 
                 50,600 
                 50,600 
                 50,600 
                 50,600 
                 50,600 
                 50,600 
               
               
                 Tail group 
                 8,800 
                 8,800 
                 8,800 
                 8,800 
                 8,800 
                 8,800 
               
               
                 Thermal protection system 
                 46,900 
                 46,900 
                 46,900 
                 46,900 
                 46,900 
                 46,900 
               
               
                 Landing gear 
                 15,100 
                 12,200 
                 14,500 
                 18,000 
                 13,300 
                 16,500 
               
               
                 Staging/separation systems 
                 2,100 
                 1,100 
                 2,500 
                 3,000 
                 1,700 
                 4,000 
               
               
                 Elect. pwr/conv/distr 
                 5,800 
                 4,500 
                 5,800 
                 6,000 
                 4,500 
                 5,800 
               
               
                 systems 
               
               
                 Surface control activation 
                 3,000 
                 3,000 
                 3,000 
                 3,000 
                 3,000 
                 3,000 
               
               
                 Avionics 
                 1,600 
                 1,600 
                 1,600 
                 1,600 
                 1,600 
                 1,600 
               
               
                 Propulsion systems (incl. 
                 18,900 
                 12,800 
                 15,900 
                 22,400 
                 18,900 
                 25,100 
               
               
                 env. control) 
               
               
                 Reaction control system 
                 2,300 
                 1,800 
                 2,200 
                 2,500 
                 2,000 
                 2,500 
               
               
                 Orbit maneuver systems 
                 0 
                 0 
                 0 
                 0 
                 0 
                 0 
               
               
                 Flight deck/life 
                 0 
                 0 
                 0 
                 0 
                 0 
                 0 
               
               
                 support/cargo bay 
               
               
                 Dry weight margin (15%) 
                 48,060 
                 39,600 
                 47,190 
                 49,600 
                 45,540 
                 49,890 
               
               
                 RD-180 engines 
                 95,240 
                 — 
                 — 
                 119,050 
                 — 
                 — 
               
               
                 SSME engines 
                 43,218 
                 101,700 
                 — 
                 43,218 
                 128,820 
                 — 
               
               
                 DF/DX engines 
                 — 
                 — 
                 113,778 
                 — 
                 — 
                 154,473 
               
               
                 Main vehicle dry weight 
                 506,918 
                 405,300 
                 475,568 
                 543,148 
                 478,260 
                 536,903 
               
               
                 Personnel 
                 0 
                 0 
                 0 
                 0 
                 0 
                 0 
               
               
                 Payload to 50 × 100 n. mi. 
                 302,000 
                 209,500 
                 373,000 
                 462,000 
                 312,000 
                 735,000 
               
               
                 orbit 
               
               
                 Reaction control propellants 
                 11,200 
                 9,100 
                 10,600 
                 11,900 
                 10,600 
                 11,900 
               
               
                 Residual propellants 
                 12,700 
                 9,000 
                 12,600 
                 12,700 
                 9,000 
                 12,600 
               
               
                 Ascent propellants 
               
               
                 Oxygen 
                 5,334,900 
                 3,867,429 
                 5,302,043 
                 5,334,900 
                 3,867,429 
                 5,302,043 
               
               
                 Hydrogen 
                 468,500 
                 644,571 
                 480,637 
                 468,500 
                 644,571 
                 480,637 
               
               
                 Kerosene (RP-1) 
                 971,100 
                 — 
                 — 
                 971,000 
                 — 
                 — 
               
               
                 Propane 
                 — 
                 — 
                 755,695 
                   
                 — 
                 755,695 
               
               
                 Drop tank inert wt., incl. 
                 — 
                 — 
                 — 
                 53,684 
                 53,000 
                 109,900 
               
               
                 residuals 
               
               
                 Drop tank propellants 
               
               
                 Oxygen 
                 — 
                 — 
                 — 
                 1,095,601 
                 1,104,099 
                 2,246,919 
               
               
                 Hydrogen 
                 — 
                 — 
                 — 
                 182,600 
                 184,016 
                 374,486 
               
               
                 Gross liftoff weight (GLOW) 
                 7,607,218 
                 5,144,900 
                 7,410,142 
                 9,136,033 
                 6,662,975 
                 10,566,093 
               
               
                   
               
             
          
         
       
     
         [0000]    
       
         
               
             
               
               
             
               
               
               
             
               
               
               
               
               
               
               
             
               
               
               
               
               
               
               
             
           
               
                 TABLE 8 
               
             
             
               
                   
               
               
                 Vehicle Mass Estimates (Main Tank Diam: 27.5 feet) 
               
             
          
           
               
                   
                 VEHICLE 
               
             
          
           
               
                   
                 FULLY REUSABLE 
                 GROWTH WITH O 2 /H 2  DROP 
               
               
                   
                 (Single Stage) 
                 TANKS (1½ - Stage) 
               
             
          
           
               
                   
                 5xRD-180 
                   
                   
                 7xRD-180 
                   
                   
               
               
                   
                 3xSSME 
                 9xSSME 
                   
                 4xSSME 
                 11xSSME 
               
               
                 WEIGHTS 
                             = 77.5 
                             = 50.0 
                 8xDF/DX 
                             = 77.5 
                             = 50.0 
                 11xDF/DX 
               
               
                   
               
             
          
           
               
                 Tankage 
                 59,200 
                 54,400 
                 57,000 
                 59,200 
                 54,400 
                 57,000 
               
               
                 Body structure 
                 44,100 
                 31,700 
                 36,600 
                 51,900 
                 38,700 
                 50,300 
               
               
                 Wing group 
                 36,000 
                 36,000 
                 36,000 
                 36,000 
                 36,000 
                 36,000 
               
               
                 Tail group 
                 6,100 
                 6,100 
                 6,100 
                 6,100 
                 6,100 
                 6,100 
               
               
                 Thermal protection system 
                 32,400 
                 32,400 
                 32,400 
                 32,400 
                 32,400 
                 32,400 
               
               
                 Landing gear 
                 10,000 
                 8,100 
                 8,800 
                 11,600 
                 8,800 
                 10,000 
               
               
                 Staging/separation systems 
                 1,200 
                 1,000 
                 1,500 
                 2,400 
                 1,500 
                 2,400 
               
               
                 Elect. pwr/conv/distr 
                 3,300 
                 2,600 
                 3,200 
                 3,300 
                 2,600 
                 3,200 
               
               
                 systems 
               
               
                 Surface control activation 
                 2,100 
                 2,100 
                 2,100 
                 2,100 
                 2,100 
                 2,100 
               
               
                 Avionics 
                 1,600 
                 1,600 
                 1,600 
                 1,600 
                 1,600 
                 1,600 
               
               
                 Propulsion systems (incl. 
                 13,100 
                 7,300 
                 9,100 
                 18,300 
                 8,900 
                 12,500 
               
               
                 env. control) 
               
               
                 Reaction control system 
                 1,500 
                 1,200 
                 1,300 
                 2,000 
                 1,300 
                 1,500 
               
               
                 Orbit maneuver systems 
                 0 
                 0 
                 0 
                 0 
                 0 
                 0 
               
               
                 Flight deck/life 
                 0 
                 0 
                 0 
                   
                   
                 0 
               
               
                 support/cargo bay 
               
               
                 Dry weight margin (15%) 
                 31,590 
                 27,675 
                 29,355 
                 34,035 
                 29,160 
                 32,265 
               
               
                 RD-180 engines 
                 59,525 
                 — 
                 — 
                 83,335 
                   
                 — 
               
               
                 SSME engines 
                 21,609 
                 61,020 
                 — 
                 28,812 
                 74,580 
                 — 
               
               
                 DF/DX engines 
                 — 
                 — 
                 65,015 
                 — 
                 — 
                 89,397 
               
               
                 Main vehicle dry weight 
                 323,324 
                 273,195 
                 290,070 
                 373,082 
                 298,140 
                 336,762 
               
               
                 Personnel 
                 0 
                 0 
                 0 
                 0 
                 0 
                 0 
               
               
                 Payload to 50 × 100 n. mi. 
                 158,000 
                 80,000 
                 197,000 
                 318,000 
                 158,000 
                 398,000 
               
               
                 orbit 
               
               
                 Reaction control propellants 
                 7,100 
                 5,800 
                 6,100 
                 8,400 
                 6,800 
                 7,200 
               
               
                 Residual propellants 
                 7,300 
                 5,700 
                 7,200 
                 7,300 
                 5,700 
                 7,600 
               
               
                 Ascent propellants 
               
               
                 Oxygen 
                 3,165,195 
                 2,222,785 
                 3,050,364 
                 3,165,195 
                 2,222,785 
                 3,050,364 
               
               
                 Hydrogen 
                 277,961 
                 370,465 
                 276,520 
                 277,961 
                 370,465 
                 276,520 
               
               
                 Kerosene (RP-1) 
                 576,093 
                 — 
                 — 
                 576,093 
                 — 
                 — 
               
               
                 Propane 
                 — 
                 — 
                 434,766 
                 — 
                 — 
                 434,766 
               
               
                 Drop tank inert wt., incl. 
                 — 
                 — 
                 — 
                 62,900 
                 32,400 
                 64,900 
               
               
                 residuals 
               
               
                 Drop tank propellants 
               
               
                 Oxygen 
                 — 
                 — 
                 — 
                 1,284,433 
                 661,456 
                 1,324,549 
               
               
                 Hydrogen 
                 — 
                 — 
                 — 
                 214,072 
                 110,243 
                 220,758 
               
               
                 Gross liftoff weight (GLOW) 
                 4,514,973 
                 2,957,925 
                 4,262,020 
                 6,287,436 
                 3,865,989 
                 6,121,419 
               
               
                   
               
             
          
         
       
     
         [0000]    
       
         
               
             
               
               
             
               
               
               
             
               
               
               
               
               
               
               
             
               
               
               
               
               
               
               
             
           
               
                 TABLE 9 
               
             
             
               
                   
               
               
                 Vehicle Mass Estimates (Main Tank Diam: 16.5 &amp; 12.5 feet) 
               
             
          
           
               
                   
                 MAIN TANK DIAMETER 
               
             
          
           
               
                   
                 16.5 feet 
                 12.5 feet 
               
               
                   
                 Delta IV Heavy Heritage 
                 Atlas V Heritage 
               
             
          
           
               
                   
                 2xRD-180 
                   
                   
                 1xRD-180 
                   
                   
               
               
                   
                 2xSSME 
                 5xSSME 
                   
                 1xSSME 
                 3xSSME 
               
               
                 WEIGHTS 
                            = 77.5 
                            = 50.0 
                 5xDF/DX 
                            = 77.5 
                            = 50.0 
                 2xDF/DX 
               
               
                   
               
             
          
           
               
                 Tankage 
                 14,825 
                 13,800 
                 14,320 
                 6,446 
                 6,002 
                 6,226 
               
               
                 Body structure 
                 17,540 
                 16,500 
                 17,720 
                 7,956 
                 6,984 
                 8,475 
               
               
                 Wing group 
                 12,650 
                 12,650 
                 12,650 
                 7,260 
                 7,260 
                 7,260 
               
               
                 Tail group 
                 2,200 
                 2,200 
                 2,200 
                 1,263 
                 1,263 
                 1,263 
               
               
                 Thermal protection system 
                 14,725 
                 14,725 
                 14,725 
                 9,500 
                 9,500 
                 9,500 
               
               
                 Landing gear 
                 3,020 
                 3,000 
                 2,900 
                 2,000 
                 1,950 
                 2,000 
               
               
                 Staging/separation systems 
                 1,000 
                 1,000 
                 1,000 
                 1,000 
                 1,000 
                 1,000 
               
               
                 Elect. pwr/conv/distr 
                 2,000 
                 2,000 
                 2,000 
                 1,500 
                 1,500 
                 1,500 
               
               
                 systems 
               
               
                 Surface control activation 
                 1,200 
                 1,200 
                 1,200 
                 1,000 
                 1,000 
                 1,000 
               
               
                 Avionics 
                 1,600 
                 1,600 
                 1,600 
                 1,600 
                 1,600 
                 1,600 
               
               
                 Propulsion systems (incl. 
                 2,560 
                 2,350 
                 2,150 
                 1,500 
                 1,380 
                 1,240 
               
               
                 env. control) 
               
               
                 Reaction control system 
                 1,100 
                 1,100 
                 1,100 
                 1,000 
                 1,000 
                 1,000 
               
               
                 Orbit maneuver systems 
                 0 
                 0 
                 0 
                 0 
                 0 
                    0 
               
               
                 Flight deck/life 
                 0 
                 0 
                 0 
                 0 
                 0 
                    0 
               
               
                 support/cargo bay 
               
               
                 Dry weight margin (15%) 
                 11,163 
                 10,819 
                 11,035 
                 6,304 
                 6,066 
                 6,310 
               
               
                 RD-180 engines 
                 23,810 
                 — 
                 — 
                 11,905 
                 — 
                 — 
               
               
                 SSME engines 
                 13,560 
                 33,900 
                 — 
                 7,203 
                 20,340 
                 — 
               
               
                 DF/DX engines 
                 — 
                 — 
                 40,635 
                 — 
                 — 
                 16,254  
               
               
                 Main vehicle dry weight 
                 122,953 
                 116,844 
                 125,235 
                 67,437 
                 66,845 
                 64,628  
               
               
                 Personnel 
                 0 
                 0 
                 0 
                 0 
                 0 
                    0 
               
               
                 Payload to 50 × 100 n. mi. 
                 98,000 
                 79,000 
                 158,000 
                 46,000 
                 27,000 
                 60,000  
               
               
                 orbit 
               
               
                 Reaction control propellants 
                 1,500 
                 1,500 
                 1,500 
                 1,000 
                 1,000 
                 1,000 
               
               
                 Residual propellants 
                 2,000 
                 2,000 
                 2,000 
                 1,500 
                 1,500 
                 1,500 
               
               
                 Ascent propellants 
               
               
                 Oxygen 
                 802,216 
                 483,214 
                 878,835 
                 349,005 
                 210,102 
                 382,118  
               
               
                 Hydrogen 
                 42,343 
                 80,536 
                 39,655 
                 18,301 
                 35,017 
                 17,242  
               
               
                 Kerosene (RP-1) 
                 210,832 
                   
                 — 
                 91,723 
                 — 
                 — 
               
               
                 Propane 
                 — 
                 — 
                 150,045 
                 — 
                 — 
                 65,240  
               
               
                 Drop tank inert wt., incl. 
                 33,900 
                 42,857 
                 47,355 
                 19,096 
                 20,590 
                 20,590  
               
               
                 residuals 
               
               
                 Drop tank propellants 
               
               
                 Oxygen 
                 692,786 
                 846,229 
                 966,643 
                 389,720 
                 420,202 
                 420,202  
               
               
                 Hydrogen 
                 115,464 
                 141,038. 
                 161,107 
                 64,953 
                 70,034 
                 70 033  
               
               
                 Gross liftoff weight (GLOW) 
                 2,121,994 
                 1,793,218 
                 2,530,375 
                 1,048,735 
                 852,290 
                 1,102,553    
               
               
                   
               
             
          
         
       
     
         [0000]    
       
         
               
             
               
               
             
               
               
               
             
               
               
               
               
               
               
               
             
               
               
               
               
               
             
           
               
                 TABLE 10 
               
               
                   
               
               
                 Summary of reusable global launcher performance in terms of 
               
               
                 payload and payload: dry mass and payload: gross mass ratios 
               
               
                   
               
             
             
               
                   
               
             
          
           
               
                   
                 VEHICLE 
               
             
          
           
               
                   
                   
                 GROWTH WITH O 2 /H 2  DROP 
               
               
                   
                 FULLY REUSABLE 
                 TANKS 
               
               
                   
                 Single Stage 
                 (1½ - Stage) 
               
             
          
           
               
                   
                 8xRD- 
                   
                   
                 10xRD- 
                   
                   
               
               
                   
                 180 
                   
                   
                 180 
               
               
                   
                 6xSSME 
                 15xSSME 
                   
                 6xSSME 
                 19xSSME 
               
               
                 WEIGHTS 
                             = 77.5 
                             = 50.0 
                 14xDF/DX 
                             = 77.5 
                             = 50.0 
                 19xDF/DX 
               
               
                   
               
               
                 33 FEET 
                 302 
                 209 
                 373 
                 462 
                 312 
                 735 
               
               
                 (SATURN 5 MAIN TANKS) 
                 0.60 
                 0.52 
                 0.78 
                 0.77 
                 0.59 
                 1.13 
               
               
                   
                 0.04 
                 0.04 
                 0.05 
                 0.05 
                 0.05 
                 0.07 
               
               
                   
               
               
                   
                 5xRD- 
               
               
                   
                 180 
                   
                   
                 7xRD-180 
               
               
                   
                 3xSSME 
                 9xSSME 
                   
                 4xSSME 
                 11xSSME 
               
               
                 (ENGINES) 
                             = 77.5 
                             = 50.0 
                 8xDF/DX 
                             = 77.5 
                             = 50.0 
                 11xDF/DX 
               
               
                   
               
               
                 27.5 FEET 
                 158 
                 80 
                 197 
                 318 
                 158 
                 398 
               
               
                 (SHUTTLE EXTERNAL 
                 0.43 
                 0.29 
                 0.73 
                 0.73 
                 0.48 
                 0.99 
               
               
                 TANK) 
                 0.03 
                 0.03 
                 0.05 
                 0.05 
                 0.04 
                 0.06 
               
               
                   
               
             
          
           
               
                   
                   
                 2xRD-180 
                   
                   
               
               
                   
                   
                 2xSSME 
                 5xSSME 
               
               
                 (ENGINES) 
                   
                             = 77.5 
                             = 50.0 
                 5xDF/DX 
               
               
                   
               
               
                 16.5 FEET 
                 MARGINAL OR 
                 98 
                 79 
                 158 
               
               
                 (DELTA IV) 
                 NOT FEASIBLE 
                 0.62 
                 0.49 
                 0.92 
               
               
                   
                   
                 0.05 
                 0.04 
                 0.06 
               
               
                   
               
               
                   
                   
                 1xRD-180 
               
               
                   
                   
                 1xSSME 
                 3xSSME 
               
               
                 (ENGINES) 
                   
                             = 77.5 
                             = 50.0 
                 2xDF/DX 
               
               
                   
               
               
                 12.5 FEET 
                 NOT FEASIBLE 
                 46 
                 27 
                 60 
               
               
                 (ATLAS V) 
                   
                 0.53 
                 0.37 
                 0.70 
               
               
                   
                   
                 0.04 
                 0.03 
                 0.05 
               
               
                   
               
               
                 Payload to 50 × 100 n.mi. orbit (10 3  lbm) 
               
               
                 Payload: Dry weight ratio 
               
               
                 Payload: Gross weight ratio 
               
             
          
         
       
     
         [0050]      FIGS. 7   a - b  shows comparisons of vehicle dry mass scaling.  FIG. 7   a  graphs comparisons of vehicle dry mass estimates of the present invention vs. vehicle total estimated volume, with dry mass estimates of Lepsch, et al. and one proposed equation from Czysz, P. A. and Rahaim, C. P., “Perspective of Launch Vehicle Size and Weight Based on Propulsion System Concept,” IAC-02-V.4.08, 53 rd  IAC, The World Space Congress, 10-19 Oct. 2002, Houston Tex. The comparison shows that the mass estimates of the present invention are 20-30 percent higher (more conservative) than those of Lepsch, et al and Czysz, et al. Accordingly, the vehicle performance estimates of the present invention are considered correspondingly conservative.  FIG. 7   b  shows a proposed reference dry mass equation relating vehicle dry mass to total vehicle volume by Czsyz et al. Metric-to-English conversion is shown. 
         [0051]      FIG. 8  shows example reusable global launcher (RGL) performance analysis, with and without an upper “kick” stage, for cargo delivery to LEO (main tank diameter 27.5 feet, 1½ stage). Results show LEO cargo increases of 28-36 percent using an existing and an advanced design “kick” stage starting from suborbital velocity and altitude, compared to direct delivery. If vehicle and flight dynamic optimization showed that “kick” stage start and/or container jettison could occur earlier in the trajectory, larger cargo increases would result. In such case, RGL powered flight could continue after staging if necessary to reach a velocity from which glide alone would achieve circum-global return and recovery. For far space missions (i.e., geosynchronous and beyond), direct delivery would be infeasible; larger “kick” stages would be required, but suborbital “kick’ stage separation and start would still confer enhanced performance. The Advanced Upper Stage is described in Slazer, Frank A., et al., “Delta IV Launch Vehicle Growth Options to Support NASA&#39;s Space Exploration Vision”, The Boeing Company, IAC-04-V.4.03, 55 th  International Astronautical Congress, Vancouver, Canada, 2004. 
         [0052]      FIG. 9  depicts use of a reusable global launcher as the Earth launch platform for a geolunar shuttle (using existing engines, with a main tank diameter of 27.5 feet). The shuttle design shown is as disclosed in IAC 05-D2.3.08, “Geolunar Shuttle as Upper Stage for Heavy Earth Launchers”, 56 th  IAC, Fukuoka, Japan, October 2005. Parameters for this embodiment are listed in Table 11. 
         [0000]    
       
         
               
               
               
             
               
               
               
               
               
             
               
               
               
               
               
             
           
               
                 TABLE 11 
               
             
             
               
                   
               
               
                   
                 REUSABLE GLOBAL LAUNCHER 
                 GEOLUNAR SHUTTLE 
               
             
          
           
               
                 VEHICLE 
                   
                 DROP  
                   
                 DROP 
               
               
                 PARAMETER 
                 CORE 
                 TANKS (2) 
                 CORE 
                 TANKS (2) 
               
               
                   
               
             
          
           
               
                 Payload capabilities a , lbm 
                 318,000 
                 — 
                 — 
                 — 
               
               
                 Crew (2), lbm 
                 = 
                   
                 500 
                 = 
               
               
                 Cargos, Earth→Moon/return, lbm 
                   
                   
                 15,000/5,000 
                   
               
               
                 Gross liftoff mas, lbm 
                 4,726,000 
                 1,561,400 
                 162,100 
                 155,900 
               
               
                 Dry mass less engines b , lbm 
                 260,900 
                 62,900 
                 34,700 
                 5,900 
               
               
                 Engines, lbm 
                   
                   
                   
                   
               
               
                 7 × RD-180 
                 83,335 
                 — 
                 — 
                 — 
               
               
                 4 × SSME (ε = 77.5) 
                 28,812 
                 — 
                 — 
                 — 
               
               
                 3 × RL10B-2 
                 — 
                 — 
                 1,992 
                 — 
               
               
                 Re-entry planform loading, lbm/ft 2   
                 30.2 
                 — 
                 16.1 
                 — 
               
               
                 Cargo bag, ft 
                 — 
                 — 
                 12 × 30 
                 — 
               
               
                 Cargo density, lb/ft 3   
                 — 
                 — 
                 4.4/1.5 
                 — 
               
               
                 Return glide downrange, n.mi. 
                 11,500 
                 — 
                 (global) 
                 — 
               
               
                 Return glide crossrange, n.mi. 
                 ±3,500 
                 — 
                 ±4,500 
                 — 
               
               
                   
               
               
                   a 50 × 100 n.mi., 28.7°;  b incl. 15% margin 
               
             
          
         
       
     
         [0053]      FIG. 10  shows an embodiment of the reusable global launcher of the present invention with subsonic air launch (using existing engines, main tank diameter 12.5 feet). The subsonic launch aircraft design is as disclosed in “Direct Flight Far Space Shuttle”, US Patent Application Publication, US2009/0140101, Jun. 4, 2009. Parameters are listed in Table 12. 
         [0000]    
       
         
               
               
               
             
               
               
               
               
             
               
               
               
               
             
           
               
                 TABLE 12 
               
             
             
               
                   
               
               
                 VEHICLE 
                   
                 REUSABLE GLOBAL LAUNCHER 
               
             
          
           
               
                 PARAMETER 
                 AIRCRAFT 
                 CORE 
                 DROP TANKS 
               
               
                   
               
             
          
           
               
                 Payload a , lbm 
                 950,000 
                 53,000 
                 — 
               
               
                 Cargo, lbm 
                 — 
                 48,100 
                 — 
               
               
                 Cargo container, lbm 
                 — 
                 4,900 
                 — 
               
               
                 Gross start mass, lbm 
                 1,896,300 
                 578,300 
                 371,100 
               
               
                 Dry mass less engines b , lbm 
                 524,040 
                 48,329 
                 14,900 
               
               
                 Engines, lbm 
                   
                   
                   
               
               
                 4 × GE-115B (less tailpipe) 
                 77,260 
                 — 
                 — 
               
               
                 2 × RD-180 
                 23,810 
                 — 
                 — 
               
               
                 1 × RD-180 
                 — 
                 11,905 
                 — 
               
               
                 1 × SSME(ε = 77.5) 
                 — 
                 7,203 
                 — 
               
               
                 Re-entry planform loading, lbm/ft 2   
                 — 
                 23.6 
                 — 
               
               
                 Cargo space, ft 
                 — 
                 12 × 70 
                 — 
               
               
                 Cargo density, lb/ft 3   
                 — 
                 6.1 
                 — 
               
               
                 Return glide downrange, n.mi. 
                 — 
                 (global) 
                 — 
               
               
                 Return glide crossrange, n.mi. 
                 — 
                 ±3,500 
                   
               
               
                   
               
               
                   a 50 × 100 n.mi., 28.7°;  b incl. 15% margin 
               
             
          
         
       
     
         [0054]      FIG. 11  shows an embodiment of a reusable global launcher with subsonic air launch as the Earth launch platform for a far space shuttle, in this far space mission shown, to perform Earth-low Moon orbit-Earth round-trip with 6 personnel and 10,000 Ibm payload (using existing engines, with main tank diameter 16.5 feet). The subsonic launch aircraft design is scaled up from that shown in  FIG. 10 , and the far space shuttle design is scaled down from that shown in  FIG. 9 . Deployment of the far space shuttle/reusable global launcher combination is accomplished using a trapeze mechanism to ensure safe separation. This embodiment preferably uses rocket assisted pullup (launch at 60,000 ft. altitude; 45° flight path angle). Parameters are listed in Table 13. 
         [0000]    
       
         
               
               
               
               
             
               
               
               
               
               
             
               
               
               
               
               
             
           
               
                 TABLE 13 
               
             
             
               
                   
               
               
                 VEHICLE 
                   
                 REUSABLE GLOBAL LAUNCHER 
                 FAR SPACE 
               
             
          
           
               
                 PARAMETER 
                 AIRCRAFT 
                 CORE 
                 DROP TANKS 
                 SHUTTLE 
               
               
                   
               
             
          
           
               
                 Nominal payload, lbm 
                 2,200,000 
                   124,000 a   
                 — 
                 — 
               
               
                 Crew 
                 ? 
                 — 
                 — 
                  2,000 (6) 
               
               
                 Cargo 
                 — 
                 — 
                 — 
                  10,000 b   
               
               
                 Gross liftoff mass, lbm 
                 4,400,000 
                 1,181,844 
                 842,150 
                 120,587 
               
               
                 Dry mass less engines, lbm 
                 1,217,000 
                   85,583 c   
                  33,900 
                  26,919 c   
               
               
                 Engines, lbm 
                   
                   
                   
                   
               
               
                 8 × GE90-115B 
                   154,520 
                 — 
                 — 
                 — 
               
               
                 3 × RD-180 
                   35,715 
                 — 
                 — 
                 — 
               
               
                 2 × RD-180 
                 — 
                   23,810 
                 — 
                 — 
               
               
                 2 × SSME(ε = 50) 
                 — 
                   13,560 
                 — 
                 — 
               
               
                 2 × RL10B-2 
                 — 
                 — 
                 — 
                  1,281 
               
               
                 Re-entry planform loading, lbm/ft 2   
                 — 
                      27.2 
                 — 
                    11.2 
               
               
                 Cargo bay, ft 
                 — 
                 — 
                 — 
                 12 × 15 
               
               
                 Cargo density, lb/ft 3   
                 — 
                 — 
                 — 
                     5.9 
               
               
                 Return glide downrange, n.mi. 
                 — 
                 (global) 
                 — 
                 (global) 
               
               
                 Return glide crossrange, n.mi. 
                 — 
                   ±3,500 
                 — 
                  ±4,500 
               
               
                   
               
               
                   a 50 × 100 n.mi., 28.7°;  b Round-Trip: Earth→ low Moon orbit→ Earth;  c incl. 15% margin 
               
             
          
         
       
     
         [0055]    Just as embodiments of the single-stage reusable global launcher (RGL) can be augmented/grown by adding external drop tanks as shown in this study, its performance can also be increased by adding propulsive “strap-on” boosters, often called “zero stages”. These can involve liquid, solid or liquid/solid hybrid propulsion. Such strap-on boost stages, with parachute or splash-down recovery, are sometimes termed “reusable”, although the amount of re-engineering required especially for salt-water immersion, renders that term questionable. 
         [0056]    Development of a reusable propulsive airframe as embodied in some RGL embodiments disclosed herein would provide a basic capability vehicle which can be grown and adapted, or “morphed”, to support and perform a range of manned and unmanned missions. For example, if it is large, its liftoff and ascent burn phase can be “human-rated” that it could serve as a reusable Earth launcher for a manned geolunar shuttle as shown in  FIG. 9 , and support for far space missions such as to Earth-sun libration points, asteroids and Mars. If it is smaller, it could serve as a ground launcher or air-launched launcher for LEO payloads as shown in  FIG. 10 . It could also be modified as a manned or unmanned single-stage-to-orbit (SSTO) shuttle. In even smaller embodiments it could be air launched from existing aircraft for various global missions. 
         [0057]    Embodiments of the present invention include but are not limited to novel combinations of one or more of the following elements: (1) simplicity including: (a) being unmanned, (b) circular cross-section tankage (separate, with no common bulkheads), (c) no internal cargo bay, and/or (d) dry wings; (2) vehicle performance advantages including: (a) dual-fuel propulsion with existing separate single-fuel engines, (b) projected dual-fuel engines; and/or (3) high hypersonic L/D (˜3) configuration. 
         [0058]    Embodiments of the present invention preferably comprise one or more of following benefits: (1) economics of reusability for all or most of the vehicle hardware; (2) circum-global glide range with aerodynamic maneuverability during ascent as well as re-entry; (3) applicability for a wide range of sizes, including both ground and air launch; (4) adaptability for vehicle performance growth by use of external drop tanks, and liquid, solid or hybrid liquid-solid propulsive “strap-on” boost stages; and/or (5) adaptability for modification to manned operation. 
         [0059]    Embodiments of the RGL can be cast as a reusable single-stage boost-glide launcher capable of circum-global glide return to airstrip recovery at its launch site or similar circum-global location. It can be considered a progenitor of eventual SSTO capability when SSTO capability can be guaranteed by real-world reductions in vehicle inert mass, accounting for dry mass, residuals, losses and contingency margins. 
         [0060]    Although the invention has been described in detail with particular reference to these embodiments, other embodiments can achieve the same results. Variations and modifications of the present invention will be obvious to those skilled in the art and it is intended to cover all such modifications and equivalents. The entire disclosures of all patents, references, and publications cited above are hereby incorporated by reference.