Figures

1 (Introduction) Andromeda Galaxy 5

2 (Introduction) MIRLIN (Mid-Infrared Large Well Imager) image of the black hole at the center of the Milky Way 6

3 (Introduction) Habitable zones of life and Earth-like solar systems 7

4 (Introduction) Two scramjet-powered space launchers for approximately Mach

12 airbreather operation 8

1.1 Spectrum of launchers/spacecraft from 1956 to 1981 16

1.2 Diameter of the Sun compared with the Moon's orbital diameter 19

1.3 Sun to near-galactic space in three segments 20

1.4 Notional round trip to space destination from Earth involving four plus and minus accelerations used to establish mission mass ratios 23

1.5 Required specific impulse as a function of spacecraft speed with some projections 27

1.6 One-way distance and travel time in Earth time 28

2.1 A look to the future space infrastructure envisioned by Boris Gubanov and Viktor Legostayev 38

2.2 A Japanese look to the future space infrastructure based on their development of an aerospace plane 39

2.3 Aerospace plane concept from Japan National Aerospace Laboratories . . . . 40

2.4 International space plans as presented to the Space Advisory Council for the Prime Minister of Japan in 1988 40

2.5 Expendable vehicles are for pioneers to open up new frontiers and establish a one-way movement of people and resources 41

2.6 Sustained use vehicle industries used to open up new economic frontiers and establish scheduled, regular, sustained two-way flows of people and resources 42

2.7 The conventional path for launcher development 43

2.8 "Soyuz" launch with "Progress" re-supply capsule 45

2.9 Proton first stage in Moscow plant 45

2.10 Energia was an approach to achieve a fully reusable, extended life launcher . 47

2.11 A model of the "Energia" showing the strap-on booster parachute packs and cylindrical payload container and the Buran space plane on the Baikonur launch complex 48

2.12 Fly-back version of the strap-on as an alternative to lifting parachutes 49

2.13 Buran after landing on its first, last, and only flight 51

2.14 Total vehicle energy approaches a constant 52

2.15 Adding the weight history shows the differentiation of the propulsion systems in terms of initial weight and the convergence to a single on-orbit value 53

2.16 The rocket advocate's vision of launchers that fly regularly to space 54

2.17 A balanced vision of launchers that fly regularly to space 54

2.18 Airbreather/rocket, single-stage-to-orbit configuration and a rocket-derived hypersonic glider, single-stage-to-orbit configuration 55

2.19 Airbreather/rocket, two-stage-to-orbit configuration with all-rocket second stage and an all-rocket hypersonic glider, two-stage-to-orbit configuration with all-rocket second stage 56

2.20 Large aircraft-based two-stage-to-orbit configuration with a combined cycle-powered waverider second stage 57

2.21 The result is that the potentials were never developed and impediments were sufficient to prevent any further hardware development of a truly sustained use space launcher 58

2.22 Our current space infrastructure, but without MIR is limited to specific LEO

and GSO without significant intra-orbit operations 59

2.23 One US look to the future space infrastructure that fully utilizes the space potential by Dr William Gaubatz when director of the McDonnell Douglas Astronautics Delta Clipper Program, circa 1999 60

2.24 Waiting time is costly for commercial space operations 61

2.25 "Bud" Redding Space Cruiser launched from a trans-atmospheric vehicle to accomplish a satellite repair 64

3.1 Comparison of payload costs to orbit, from 1971 to 2003 70

3.2 Payload costs per pound based on fleet flight rate 72

3.3 Weight ratio to achieve a 100 nautical mile orbit decrease as maximum airbreathing Mach number increases 74

3.4 The less oxidizer carried, the lower the mass ratio 74

3.5 Orbital velocity decreases as altitude increases 76

3.6 Slower orbital speed means longer periods of rotation 77

3.7 To achieve a higher orbit requires additional propellant 77

3.8 Space and atmospheric vehicle development converge 80

3.9 Controlling drag 81

3.10 Wetted area parameter from Figure 3.9 correlates with Kiichemann's tau yielding a geometric relationship to describe the delta planform configurations of different cross-sectional shape 83

3.11 Hypersonic rocket-powered glider for airbreathing Mach <6 and hypersonic combined cycle-powered aircraft for airbreathing Mach >6 84

3.12 Wind-tunnel model configurations for tail effectiveness determination over hypersonic to subsonic speed regime 85

3.13 BOR V after return from hypersonic test flight at Mach 22 86

3.14 FDL-7 C/D compared with Model 176 87

3.15 Model 176 in the McDonnell Douglas Hypervelocity Impulse Tunnel (circa

3.16 FDL-7 C/D, Model 176 entry temperature distribution 89

3.17 FDL-7 C/D, Model 176 materials, thermal protection systems distribution based on temperature profile in Figure 3.16 90

3.18 McDonnell Aircraft Company Roll-Bonded Titanium Structure (circa 1963) 90

3.19 USAF one-half-scale FDL-5 vehicle 91

3.20 Individual Model 176 launch costs for a 100-launch program, as projected in a McDonnell Douglas Astronautics Corporation 1964 brief 92

3.21 USAF FDL-7C as configured by McDonnell Douglas with an escape module capable of controlled hypersonic flight 93

3.22 USAF FDL-7C/Model 176 equipped with a switchblade wing and retractable inward-turning inlet for airbreathing rocket applications 94

3.23 Takeoff and landing speeds of minimum-sized launchers 95

3.24 Imposed horizontal takeoff requirement can radically increase takeoff gross weight unless the weight ratio is less than 4.5 96

3.25 Size-determining parameter group correlates with Kiichemann's tau 98

3.26 All-rocket available design space is limited 99

3.27 The Mach 8 combined cycle launcher is also limited 100

3.28 The Mach 12 combined cycle launcher is also limited 101

3.29 Combined cycle propulsion has the advantage 102

4.1 Comparison of XLR-29 qualification (circa 1965) with that of the Space Shuttle main engine (SSME) (circa 1972) 108

4.2 Liquid rocket engine carries its fuel and oxidizer onboard 109

4.3 Airflow energy compared with available chemical energy 111

4.4 Four representative ram/scramjet module configurations 115

4.5 Four very different internal drags for the four module configurations 116

4.6 Module configuration significantly affects performance 118

4.7 Operating boundaries of Brayton cycle engines based on enthalpy and entropy analyses 120

4.8A Performance envelope of six materials 5 feet (1.52 m) aft of the nose on a full-

size operational vehicle 123

4.8B Detail performance envelope for 1,700°F (927°C) and 2,100°F (1,147°C)

material 123

4.9 Materials and engine operating regimes compared 127

4.10 Rocket-derived propulsion 133

4.11 Airbreathing rockets 136

4.12 Variable capture area, inward-turning inlet 137

4.13 Airbreathing rocket configuration concept 138

4.14 KLIN cycle, thermally integrated turbojet rocket 139

4.15 Airbreathing rocket thermally integrated combined cycle 140

4.16 Benefits of thermal integration 141

4.17 System thermal integration 142

4.18 Closed cycle heat pump and combustor fuel injection 143

4.19 System thermal integrated specific impulse 144

4.20 Integrated ejector ram-scramjet rocket 146

4.21 300°C hydrogen injected into supersonic air stream at flight conditions corresponding to a scramjet combustor for an aircraft flying at Mach 8 . . . . color

4.22 Air collection and enrichment cycle 149

4.23 The less the weight ratio, the less the oxidizer carried 150

4.24 The pulse detonation rocket engine operational cycle color

4.25 The pulse detonation engine cycle compared with the Brayton cycle 161

4.26 Pulse detonation rocket engine 162

4.27 Integrated PDRE ramjet combined cycle 163

4.28 Integrated PDRE ram-scramjet combined cycle 164

4.29 The PDE improves the total weight ratio 166

4.30 Engine thrust-to-weight ratio decreases with weight ratio 169

4.31 Gross weight decreases significantly as oxidizer-to-fuel ratio decreases 174

4.32 Gross weight decreases significantly as weight ratio decreases 175

4.33 Total volume decreases as the weight ratio decreases, except for ACES propulsion system 177

4.34 Empty weight is less if total volume is less 177

4.35 LACE rocket-powered VTOHL SSTO 179

4.36 Ejector ram-scramjet powered HTOL SSTO 180

4.37 Two elegant TSTO designs 181

4.38 Comparison of SSTO and TSTO results for TOGW 183

4.39 Comparison of SSTO and TSTO results for OEW 183

4.40 Ajax 187

4.41 Ayaks illustration by Alexandre Szames from information obtained from Vladimir Freistat, the Program Director of AYAKS 189

4.42 Laser/microwave-heated MHD spacecraft operating envelope enabled by a series of propulsion configuration adaptations 193

4.43 Laser/microwave-heated MHD spacecraft 194

4.44 Sketch of variable cycle ramjet based on Rocketdyne SSME, circa 1983. . . . 195

4.45 Alternate expansion nozzle configuration 196

5.1 Future space infrastructure by Dr William Gaubatz 210

5.2 Launch velocity increment to reach Earth orbit 214

5.3 Velocity increment to 200 nautical mile orbit for orbital inclination 215

5.4 Launch propellant required to lift orbital maneuver propellant to LEO by a rocket ejector ramjet 218

5.5 Propellant required parametrics with respect to payload mass and density . . 218

5.6 Transfer ellipse to change orbital altitude 221

5.7 Velocity requirement to change orbital altitude can approach one-half of the orbital speed 224

5.8 Mass ratio required to change orbital altitude is very dependent on propulsion system performance 225

5.9 Ratio of total propellant weight/satellite weight 228

5.10 Ratio of total propellant weight/satellite weight for nuclear electric propulsion 229

5.11 Orbital plane change via an aerodynamic turn in the upper atmosphere and an impulse turn executed during an elliptical transfer orbit to 22,400 nautical mile orbit 231

5.12 Velocity increment to rotate orbital plane for different orbital altitudes 232

5.13 Velocity increment as a function of turn method 233

5.14 Aerodynamic turn at 245,000 ft at 22,000 ft/s 234

5.15 Mass ratio requirements for orbital plane change 235

5.16 Ratio of total propellant weight to satellite weight 238

5.17 Ratio of total propellant weight to satellite weight for solar and nuclear electric propulsion 238

5.18 Relative size and general configuration of OMVs 241

5.19 LEO-GSO-LEO two-way OMV with shield 241

5.20 OMV for impulse turn and hypersonic glider for aerodynamic turn 242

5.21 Orbital maneuver missions per 191 propellant payload for five different OMV propulsion systems 243

5.22 Large orbital station in final assembly and integration with its PROTON launcher 245

5.23 Student design team results for requirements in terms of orbital systems hardware 246

5.24 An orbital infrastructure station fabricated from discarded Shuttle main propellant tanks with a Space Shuttle docked for resupply 248

5.25 An orbital infrastructure station fabricated from discarded Shuttle main propellant tanks with a hypersonic glider resupply spacecraft analogous to MDC model 176 249

5.26 ''Bud'' Redding Space Cruiser launched from a trans-atmospheric vehicle to accomplish a satellite repair 251

5.27 An orbital infrastructure station fabricated from discarded Shuttle main propellant tanks with docked In-Space Operations Corporation Space Cruiser, a hypersonic orbital plane change vehicle and OMVs 251

6.1 A Presidential Study to continue exploration in the future color

6.2 Orbital parameters of the Moon and distances from Earth 256

6.3 The Earth-Moon system revolves about the barycenter some 4,600 km from the center of the Earth 257

6.4 Flight path geometry of the representative lunar trajectory 258

6.5 Earth orbit injection speed is less than escape speed, so the trajectory to the Moon is a transfer ellipse analogous to LEO to GSO transfer ellipse 259

6.6 Transfer trajectory from Earth orbit to lunar orbit from a brief by V. Gubanov at the European Space Conference in Bonn, Germany, in 1984 260

6.7 Superconducing MagLev launcher on the Moon to provide a non-chemical propulsion means to achieve lunar escape speed 264

6.8 We have been there before with probes, landers, orbiters, and human visitors 269

6.9 Orbital station MIR in its 15th and last year of operation 272

6.10 International Space Station in orbit 273

6.11 ESA concept for underground lunar habitat 275

6.12 ESA concept for long-term lunar structures 275

6.13 From Thomas Stafford's Report to Congress: comparison of representative lunar sites with representative Martian sites 277

6.14 The far side of the Moon from Soviet Luna 3 spacecraft compared with the near side 279

6.15 Moon topography from the laser ranger measurements by Clementine and Lunar Prospector spacecraft color

6.16 Photo of Earth-rise from Apollo 10 command module in lunar orbit color

7.1 Minimum DV to reach selected destinations in our Solar System 284

7.2 Features and average distances of objects from the Sun 284

7.3 Increased Isp reduces transit time and weight ratio 287

7.4 Comparison between chemical and nuclear sources 290

7.5 Structure and size of a NERVA-type fuel bar 290

7.6 Velocity gained by leftover fuel mass in fission as a function of percentage a of mass fissioned 295

7.7 Conceptual scheme of a nuclear thermal rocket 298

7.8 Conceptual scheme of a nuclear electric rocket 299

7.9 Types of radiation emitted from a fission reactor 301

7.10 Gamma-ray absorption coefficient for some materials 303

7.11 Absorption coefficient p and p/p of 4MeV gamma-rays in some materials. . 304

7.12 5 MeV neutron and gamma-ray relaxation length for some materials 306

7.13 Diagram of a NERVA Kiwi nuclear reactor showing a single fuel bar cross-section 308

7.14 The NERVA Kiwi B4-E reactor on its test stand at Los Alamos 309

7.15 The 4GW PHOEBUS-2 nuclear reactor on its test stand at Los Alamos ... 310

7.16 Schematic diagram of the Westinghouse NRX nuclear engine 310

7.17 Mock-up of the NERVA-1 as it stands in Huntsville, Alabama, Space Park 312

7.18 Simple scheme of a nuclear thermal rocket fed with liquid hydrogen 312

7.19 Cosmic ray energy spectrum 321

7.20 Westinghouse NRX XE experimental nuclear engine on its test stand 326

7.21 Schematic drawing of a particle bed reactor with power controlled by a rotating drum 328

7.22 Fuel element structure and assembly inside a MITEE reactor 330

7.23 Comparison between propulsion systems, including a high-temperature MITEE rocket, for interplanetary missions 331

7.24 Gas core reactor: schematic operation of an open cycle 333

7.25 Closed-cycle gas core reactor (conceptual scheme) 333

7.26 Diagram of a generic FF-heated Rubbia's engine 337

7.27 A conceptual scheme of the operation of a thin filament 338

7.28 Artist's view of a filament fission-powered spacecraft 338

7.29(a) Acceleration time for a spacecraft of mass 10,000 kg and 100,000 kg as a function of power P and Isp 345

7.29(b) Propellant mass for a spacecraft of mass 10,000 kg and 100,000 kg as a function of power P and Isp 346

7.29(c) A V for a spacecraft of mass 10,000 kg and 100,000 kg as a function of power P

and Isp 346

7.30 The thrust vs. Isp dilemma at fixed power 350

7.31 Generic hybrid nuclear thermal and nuclear electric rocket 352

7.32 Schematic of variable specific impulse magnetoplasma rocket concept 355

7.33 Thrust and propellant rate vs. specific impulse 356

7.34 30-day spiral trajectory from Earth and transfer to Mars 356

7.35 7-day spiral trajectory from Mars and return to Earth using a VASIMR . . . 357

7.36 Schematic view of the system for a VASIMR flight experiment 358

7.37 VASIMR technology development roadmap 359

7.38 Simplified scheme of hybrid nuclear thermal and chemical (LANTR) engine 360

8.1 The nearest stars 376

8.2 The Sun bending light acts as a lens 380

8.3 Chemical, fission, and fusion energy release and their mass conversion fractions 383

8.4 Power and Isp of chemical, fusion and fission system 386

8.5 An artist's view of a future heavy-lift vehicle in LEO 389

8.6 Binding energy per nucleon, as a function of mass number 391

8.7 Sketch of D-T fusion process 392

8.8 Fusion kinetics 396

8.9 Schematic illustration of a mirror MCF rocket 405

8.10 Tokamak geometry and magnets 406

8.11 Plasma current and B field lines in a tokamak 407

8.12 Schematic view of a tokamak MCF rocket using a divertor 407

8.13 Schematic of an advanced (spherical torus) tokamak reactor (spheromak) showing first wall and thermal insulation 408

8.14 Schematics of a dense plasma focus reactor and of a rocket operating according to its principle 409

8.15 Conceptual design of a shield system for a tokamak reactor, including the lithium cooling system breeding tritium 410

8.16 Schematic operation of inertial confinement fusion 414

8.17 Conceptual operation of an inertial confinement fusion reactor rocket with its magnetic nozzle 415

8.18 Sketch of MICF pellet 417

8.19 Conceptual scheme of an inertial electrostatic confinement reactor and of the radial distribution of its electric acceleration potential 420

8.20 Sketch of an RFC reactor with neutral beam port 422

8.21 Mass budgets for two MCF propulsion systems 424

8.22 ICF propulsion system—mass budget 425

8.23 Schematic view of the VISTA ICF rocket-powered vehicle 426

9.1 Andromeda Galaxy 438

9.2 Journey time as a function of spacecraft speed 438

9.3 Specific examples of Earth vs. ship times 442

9.4 Flight profile and differences between crew and Earth times 443

9.5 What time is on Mars? 445

9.6 The ship jumps out of conventional space into Einstein space-time 455

9.7 High acceleration results in shorter Earth trip times 456

A.1 Weighting factors for different types of radiation 467

A.2 Weighting factors for neutrons 467

A.3 Weighting factors for tissues/organs 468

A.4 Threshold for deterministic effects 470

A.5 Excessive relative risk at 1 Sv 472

A.6 Excessive absolute risk at 1 Sv 472

A.7 Uranium-238 decay chain 474

A.8 Thorium-232 decay chain 475

A.9 Uranium-235 decay chain 475

A.10 Mean dose value for natural background radiation 476

A.11 Average dose from medical use 477

A.12 Doses from some examinations 477

A.13 Number of weapons tests per year 478

A.14 Doses from weapons tests 478

A.15 Annual pro capite doses in the year 2000 480

A.16 Comparison of doses from different sources 484

B.1 Payload fraction vs. velocity ratio 489

B.2 Fusion Maxwellian reactivity 495

B.3 Lawson criterion 496

B.4 Generic fusion rocket geometry 496

B.5 Idealized power flow in a fusion rocket 497

B.6 Simple mirror field configurations 506

B.7 Baseball coils 508

B.8 Tandem mirror 509

B.9 Axial profiles in a tandem mirror 509

B.10 Schematic view of the GAMMA 10 tandem mirror 512

B.11 Field-reversed mirror 513

B.12 Layout of a gasdynamic mirror 514

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