Orbital Transfer Vehicle Research Articles

Solar thermal OTV—Applications to reusable and expendable launch vehicles

The Solar Orbit Transfer Vehicle (SOTV) program being sponsored by the U.S. Air Force Research Laboratory (AFRL) is developing technology that will engender revolutionary benefits to satellites and orbitto-orbit transfer systems. Solar thermal propulsion offers significant advantages for near-term expendable launch vehicles (ELVs) such as Delta IV, mid- to farterm reusable launch vehicles (RLVs) and ultimately to manned exploration of the Moon and Mars. Solar thermal propulsion uses a relatively large mirrored concentrator to focus solar energy onto a compact absorber, which is in turn heated to > 2200 K. This heat can then be used in two major ways. By flowing hydrogen or another working fluid through the absorber, high efficiency thrust can be generated with 800 sec or more specific impulse (Isp), almost twice that of conventional cryogenic stages and comparable with typical solid-core nuclear thermal stages. Within a decade, advances in materials and fabrication processes hold the promise of the Isp ranging up to 1,100 sec. In addition, attached thermionic or alkali metal thermoelectric converter (AMTEC) power converters can be used to generate 20 to 100 kilowatts (kW) of electricity. The SOTV Space Experiment (SOTV-SE), planned to be flown in 2003, will demonstrate both hydrogen propulsion and thermionic power generation, including advanced lightweight deployable concentrators suitable for large-scale applications. Evolutionary geosynchronous-transfer orbit/ geosynchronous-Earth orbit (GTO/GEO) payload lift capability improvements of 50% or more to the Delta IV launch vehicles could be implemented as part of the Delta IV P4I plan shortly thereafter. Beyond that, SOTV technology should allow long-term storage of stages in orbits up to GEO with tremendous maneuvering capability, potentially 4 to 5 km/sec or more. Servicing of low-Earth orbit (LEO) and GEO assets and reusable (ROTVs) are other possible applications. Offering a combination of high Isp and high thrust/weight together with thrust levels from 10 to 100 lb or more, a large-scale SOTV offers a low-cost alternative to nuclear thermal or solar electric stages for cargo missions to the Moon and Mars. This paper describes the SOTV-SE as well as these and other potential applications of this integrated power/propulsion technology.

Transient response of gallium arsenide and silicon solar cells under laser pulse

Solar cells can be used as receivers in laser power beaming applications [1]. Power beamed to a photovoltaic array could power a satellite, an orbital transfer vehicle or a lunar base. Understanding solar cell response to the pulsed output of a free-electron laser (FEL) is important for evaluation of power-beaming applications. In this work we investigate the time response of gallium arsenide and silicon solar cells to a 25 ns monochromatic pulse input. The PC-1D computer code is used to analyze the cell current during and after the pulse for various conditions. For GaAs cells, current decay was studied at 511, 840 and 870 nm. Most of the results have been calculated for a peak intensity of 50 W cm −2, which corresponds to roughly 500 suns peak concentration. Silicon cells have a longer minority carrier lifetime and weaker optical absorption, resulting in longer characteristic time constants. Si cell performance was studied at 900, 950 and 1060 nm wavelengths.

A method for solar sailing in a low earth orbit

It has been widely accepted that solar sails are incapable of operating in low earth orbits due to drag from the residual atmosphere overcoming the sunlight induced thrust. This limitation seems to require that a solar sail always maintain an altitude greater than approximately 1000 km above the earth’s surface. Placing a sail at this altitude first requires a fairly large launch vehicle which increases the cost of any proposed solar sail mission or proof of concept flight. The inability to access low altitudes also has effectively negated the possibility of using a solar sail as a low cost orbital transfer vehicle. This paper describes a new configuration of solar sail that will allow operation in high inclination earth orbits from under 350 km circular out to interplanetary space. Two possible sail configurations are described and their operation, advantages over previous sail designs and limitations are discussed. Spacecraft drag reduction technique involving the use of a shield made of a solar sail like material is developed.

Reusable Orbital Transfer Vehicle Applications of an Integrated Solar Upper Stage

There has been resurgent interest in the development of a reusable orbit transfer vehicle, based in lowEarth-orbit, for satellite placement, repositioning, recovery, and control. An integrated solar upper-stage system is examined for several near-Earth applications, including low-Earth to geosynchronous to lowEarth orbit ferrying. An integrated solar upper stage (a solar thermal power and propulsion system) permits rapid access to high orbits because of its moderate thrust levels (tens to hundreds of pounds) at specific impulses approaching 900 s, intermediate between high-performance chemical propulsion and electrical propulsion systems. An integrated solar upper-stage reusable orbit transfer vehicle can be used in two novel modes: 1) Drag return, in which large, rigid solar concentrators provide an aeroassist capability lowering total delta-V requirements; and 2) a hybridized integrated solar upper stage, coupled with and providing power to a xenon ion thruster, allowing more efficient maneuvers in nontime critical applications.

Low-Thrust Eccentricity-Constrained Orbit Raising

Covers advancements in spacecraft and tactical and strategic missile systems, including subsystem design and application, mission design and analysis, materials and structures, developments in space sciences, space processing and manufacturing, space operations, and applications of space technologies to other fields.

Electron-vibration energy exchange models in nitrogen plasma flows

This work presents an examination of the validity of the simple linear Landau-Teller-type model proposed by Lee for the electron-vibration energy exchange term in nitrogen [in Thermal Design of Aeroassisted Orbital Transfer Vehicles, edited by H. F. Nelson (AIAA, New York, 1985), Vol. 96, p. 3; in Thermophysical Aspects of Re-entry Flows, edited by J. N. Moss and C. D. Scott (AIAA, New York, 1986), Vol. 103, p. 197]. Plasma flow conditions encountered in high enthalpy wind tunnels are considered. The time-dependent relaxation of the vibrational energy of nitrogen due to electron inelastic collisions is calculated. The influence of the anharmonicity of the molecule and of the initial vibrational temperature ${\mathrm{T}}_{\mathrm{v}}$ is studied. With a harmonic oscillator approximation, it is found that a linear Landau-Teller-type model is accurate to describe the vibrational energy relaxation rate for electron temperatures ${\mathrm{T}}_{\mathrm{e}}$ in the range 3000 K\ensuremath{\leqslant}${\mathrm{T}}_{\mathrm{e}}$\ensuremath{\leqslant}20 000 K. When ${\mathrm{T}}_{\mathrm{v}}$${\mathrm{T}}_{\mathrm{e}}$, our results show that the initial model proposed by Lee overestimates by a mean factor of 3 the electron-vibration (e-V) relaxation time. When ${\mathrm{T}}_{\mathrm{v}}$>${\mathrm{T}}_{\mathrm{e}}$, the relaxation time appears to depend on the initial vibrational distribution. When the anharmonicity of nitrogen is taken into account, generally the relaxation rate of the vibrational energy deviates from a linear rate equation. In this case, it appears much more difficult to model accurately the e-V energy exchange term.

New control problems associated with a proposed future space transportation infrastructure

Operation management is an essential technology for the “OTV (Orbital Transfer Vehicle) Network” which has been proposed as a promising space transportation infrastructure for the next century. In addition to the nominal mode scheduling, the management system must quickly re-make the schedule in the case of failures of some system elements, inputs of urgent missions or mission alterations. This paper analyses this management as a new control problem, and proposes a distributed artificial intelligence-type system architecture to deal with it. The effectiveness of the system has been verified by simulation studies assuming practical problem settings.

Robust predictor-corrector guidance for aeroassisted orbital transfer

The safety zone guidance, relative to the exit phase of the atmospheric pass of an aeroassisted orbital transfer (AOT) vehicle is extended, so as to incorporate both the entry and exit phases of the atmospheric pass and to include the automatic determination of the switch time from entry to exit guidance. The guidance goals are 1) to contain the peak heating rate during the atmospheric pass, 2) to ensure that the AOT vehicle exits the atmosphere in the correct orbital plane, ascending afterward to the specified apogee, and 3) to achieve robustness vis-a-vis severe parameter dispersion effects resulting from navigation errors, atmospheric density errors, aerodynamic coefficient errors, and system errors. Numerical tests show that this guidance scheme exhibits high performance and relatively high robustness vis-a-vis dispersion effects resulting from these errors. The guidance robustness is due to two factors: 1) the switch time from entry phase to exit phase is the result of a proper engineering compromise and 2) for the exit phase, the control is not determined by aiming only at the desired apogee; instead, for each altitude-velocity pair, a special value of the path inclination is determined by maximizing the control margin while aiming at the desired apogee during the atmospheric pass.

Near-optimal highly robust guidance for aeroassisted orbital transfer

One of the difficulties in aeroassisted orbital transfer vehicle design is the complexity of the guidance scheme. In present aeroassisted orbital transfer guidance schemes, the reference value of the bank angle or the climb rate is determined via a predictor-corrector algorithm so that the desired apogee can be reached after the vehicle exits the atmosphere. These guidance schemes are quite sensitive to parameter dispersion effects due to atmospheric density errors, system errors, navigation errors, and aerodynamic coefficient errors. An innovative guidance scheme, called safety zone guidance, is developed. In this highly robust guidance, the control is not determined by aiming only at the desired apogee; instead, for each altitude-velocity pair, a special value of the path inclination is determined by maximizing the control margin while aiming at the desired apogee during the atmospheric pass. This results in a robust and accurate guidance scheme as shown by extensive numerical tests.

Status report on the U.S. space nuclear program

During the 1980s and 1990s the U.S. sponsored a number of space nuclear power and propulsion programs and studies covering many different mission plans. These space nuclear activities spanned the range from the 300-watt radioisotope thermoelectric generators currently in use on the Galileo mission to Jupiter and the Ulysses mission to explore the polar regions of the Sun to multi-megawatt space nuclear reactor concepts that could be used to provide power for neutral particle beams, free electron lasers, electromagnetic launchers, and orbital transfer vehicles. Studies were undertaken on both nuclear electric propulsion (NEP) and nuclear thermal propulsion (NTP) largely in support the proposed Space Exploration Initiative (SEI). Upgrades to radioisotope power sources involving dynamic (turbine-alternator and linear oscillator) and improved static (thermophotovoltaic and alkali-metal thermal-to-electric conversion) conversion systems were investigated. A major undertaking in the 1980s and early 1990s was the SP-100 space nuclear reactor power system which was designed to span the power range from 10 kWe to 1000 kWe to encompass most high-power missions then being studied. In addition, various efforts to explore the uses of thermionic reactors were undertaken, including an initiative with Russia to test the ENISY (TOPAZ II) thermionic reactor. This paper summarizes these activities and highlights future trends in an era of smaller spacecraft that will be built and launched in shorter periods of time under more tightly constrained budgets.

イオン推進機の軌道間輸送機への適用検討

It is expected that an orbit transfer vehicle using an ion thruster (IOTV) can obtain a large payload ratio, because the ion thruster has a high specific impulse (several thousands seconds). However, orbit transfer missions by means of the IOTV require large trip times, therefore it is necessary for the ion thruster to have a sufficient endurance performance. In this study, the trajectory of the solar powered IOTV was analyzed in consideration of the influence of earth's eclipse, then the payload capability of the IOTV and the steady/endurance performance of the ion thruster were evaluated. For results, it was found that the IOTV could achieve approximately double the payload mass of the OTV using a chemical rocket, and a required durability to the ion thruster was comparable to one that has used for ETS-VI.

Numerical Simulation of Hypersonic Non-Equilibrium Wake Flow Using a Higher-Resolution Method.

An object which flies hypersonic speed, such as an aeroassisted orbital transfer vehicle, is affected greatly by not only the bow shock but also the expansion around the shoulder and the recompression of separated flow. Though the experiments to analyze this flow have been performed by some groups recently, there are some difficulties in measuring the ow because of limitations of instrumentation. Both experimental and numerical approaches are necessary to analyze the phenomena. The efficient numerical code which has been developed by the authors for hypersonic thermochemical non-equilibrium flow is applied to simulate the problem. The calculated results of different accuracy's in space, the perfect gas and the experimental data are compared.

Neighboring optimal trajectories for aeroassisted orbital transfer under uncertainties

A neighboring optimal guidance scheme is presented for a nonlinear dynamic system in the presence of modeling and measurement uncertainties as applicable to fuel optimal control of an aeroassisted orbital transfer vehicle (AOTV). Here the vehicle is expected to follow a desired reference path with minimum error. The measurements considered include tracking of altitude and altitude rate. The solution approach is based on nonlinear control optimization, deterministic perturbation, stochastic state estimation, and linearized stochastic control that leads to a closed-loop control law. The reference state trajectory and control trajectory are precomputed by solving a two-point boundary-value control problem. Numerical results show that the actual (true) state trajectory is kept near its ideal (reference) state trajectory at the expense of increasing the total cost.

Lunar oxygen: Economic benefit for transportation systems and in situ production by fluorination

The use of oxygen produced on the Moon—called “MOONLOX”—is considered as a propellant component for a reusable Earth-Moon transportation system consisting of an aeroassisted orbital transfer vehicle and a lunar bus for lunar descent/ascent. Conditions for economic benefit are discussed and the processing concept of a lunar oxygen plant based on fluorination is presented. It is shown that the necessary mass of supply from Earth for MOONLOX-production is an important parameter, which may not be neglected due to its strong influence on the economy. In the ideal case where no supplies from Earth are required a reduction of up to 50% in masses to be launched into low Earth orbit can be obtained for a typical lunar mission with use of MOONLOX compared to a reference scenario with Earth-derived propellant. Mass-saving decreases, however, significantly with increasing supply from Earth until a critical supply-rate is reached—measured in percentage of MOONLOX-mass produced and consumed—beyond which mass-saving and thus economically promising lunar oxygen production is no longer possible. This critical supply-rate depends on the scenario for MOONLOX-utilization and is much larger in the case of in situ use of MOONLOX on the lunar surface, e.g. as ascent propellant for the lunar bus, than in the case of export for complete refuelling of both space vehicles. The latter scenario therefore requires significantly more autonomy for MOONLOX-production. The reduction of masses to be transported into low Earth orbit and corresponding MOONLOX-consumption define for given specific Earth-to-LEO transportation costs an upper limit on MOONLOX-production costs beyond which economic benefit is not possible. Depending on the MOONLOX-utilization strategy this upper limit varies between 3000 and 55000 $/kg for current Earth-to-LEO transportation costs.

Neighboring optimal guidance for aeroassisted orbital transfer

A neighboring optimal guidance scheme for a nonlinear dynamic system is devised with stochastic inputs and perfect measurements as applicable to fuel optimal control of an aeroassisted orbital transfer vehicle. For the deterministic nonlinear dynamic system describing the atmospheric maneuver, a nominal trajectory is determined. Then, taking modeling uncertainties into account, a linear, stochastic, neighboring optimal guidance scheme is devised. Assuming the additive character of the stochastic effects, the optimal trajectory is approximated as the sum of the deterministic nominal trajectory and the stochastic neighboring optimal solution. Numerical results are presented for a typical aeroassisted orbital transfer vehicle.< <ETX xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink">&gt;</ETX>

Electric orbit transfer vehicle cryogenic propellant system

An electric orbit transfer vehicle (EOTV) is intended to transfer payloads from low Earth orbit (LEO) to higher orbits using low-thrust solar-electric propulsion and hydrogen propellant. Because of its high specific impulse and synergistic sharing of power supply, attitude control and communication systems with the payload, the highly efficient EOTV transfer stage permits use of a smaller, less costly launch vehicle than if orbit transfer were accomplished using chemical propulsion. Study of the propellant storage and supply system for an EOTV intended to fly a 168 day spiral trajectory from LEO to geosynchronous orbit (GEO) reveals that the low propellant flow rate needed by the thrusters can be supplied by the boil-off from the storage tank, eliminating the need for any overboard venting. The tank can be fabricated under the same pressure-stabilized, thin, stainless steel monocoque construction as the current Centaur upper stage, and insulated with Centaur fixed foam and MLI. The tank contains a thermodynamic vent system (TVS) for control of tank pressure in zero and low gravity and for supply of propellant to the thrusters. An external compressor, accumulator and regulator condition the hydrogen boil-off provided by the TVS and provide for start-up and shut-down transients. The resulting system is simple, has a very low structural mass fraction and builds on the Centaur cryogenic upper stage technology, which has been operational for over 25 years.

Measurement of shock-wave/boundary-layer interaction in a free-piston shock tunnel

The short test times available in free-piston shock tunnels necessitate careful examination of experimental techniques. Measurements of shock-wave/bo undary-layer interaction are presented to demonstrate the effec- tiveness of these hypervelocity facilities in studies of a class of flows relevant to scramjet inlets. A two-dimen- sional laminar boundary layer with a freestream Mach number of 5.6, a static pressure of 4 kPa, and a unit Reynolds number of 1.14x 106/m was used as the incoming flow condition on a cold flat plate. A shock wave was generated by a wedge normal to the plate and with a face at 2.84 deg to the freestream direction. It is found that, for this weak interaction, steady flow conditions can be obtained, although the test time is of the order of only 1 ms. HE present high level of interest in aerospace planes and orbital transfer vehicles has focused attention on a num- ber of hypersonic fluid flow problems. One such problem is the unavoidable interaction between shock waves and boundary layers, both in the external flow over a hypersonic vehicle and in the inlet to an air-breathing propulsion system. Shock-wave/boundary-layer interaction is extremely impor- tant in the inlet of the air frame integrated scramjet engine under consideration for the next generation of space launch vehicle. The shock waves that are necessary to compress the inlet flow must glance across the boundary layer developed on the forebody of the vehicle. Similar interactions will occur farther inside the scramjet. The sudden pressure rise through a shock wave causes thickening of the boundary layer in the region of the shock wave and can cause the flow to separate. In closed ducts such as scramjet inlets the extreme result is choking of the flow, with a shock wave standing outside the entrance to the inlet and flow spilling around the sides. Inter- actions can also increase the rate of shear locally in the flow, with hot spots developing on surfaces. These phenomena in scramjet engines are examples of an important class of shock-wave/boundary-layer interactions in which sweepback of the shock wave is a key element. The flow in such interactions is then three dimensional. The configura- tion of interest in this paper is shown in Fig. 1. A planar shock wave generated by a wedge is swept back across a two-dimen- sional boundary layer on a flat plate. The leading edge of the wedge is normal to the plate. If the deflection angle of the wedge is small (less a few degrees), large regions of separated boundary layer generally do not occur and the interaction is regarded as weak. This three-dimensional interaction has been studied extensively13 at freestream Mach numbers of about 3, but investigators disagree over some basic aspects of the flow. There is a dearth of experimental data at hypersonic Mach numbers, particularly at test conditions similar to those expe- rienced in upper atmospheric powered flight and in re-entry.

Historical development of glycidyl azide polymer

The glycidyl azide polymer (GAP) is an energetic, thermally stable, insensitive, hydroxyl-terminated prepolymer with potential applications in advanced solid propellants including 1) class 1.3 nondetonable minimum smoke propellants, 2) clean propellants for solid rocket boosters, 3) gas generators/aircraft starter cartridges, 4) low-cost ASAT maneuver propulsion systems, and 5) high-performance space propellants for orbital transfer vehicles. The historical development of GAP from its conception is reviewed in this paper.

Axial compression corner flow with shock impingement

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ＡＯＴＶの大気圏飛行における最適軌道について

Many studies have shown the advantages of an Aeroassisted Orbital Transfer Vehicle (AOTV) which uses aerodynamic forces during an atmospheric pass to achieve orbital changes. This paper discusses optimal aeroassisted maneuvers to return from Geosynchronous Earth Orbit (GEO) to Low Earth Orbit (LEO). A trajectory optimization problem minimizing the total velocity change was formulated as a nonlinear programming problem, and numerical solutions were computed by the Sequential Quadratic Programming (SQP) method. The optimal trajectories are useful for the performance evaluation of vehicles or the design of flight profiles.