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At least 235 records · Page 13
All sourcesNASA NTRSDOE OSTI
NASA NTRS2021

Conditions in the WR 140 Wind-collision Region Revealed by the 1.083-um He I Line Profile

We present spectroscopy of the P Cygni profile of the 1.083-μm He i line in the WC7 + O5 colliding-wind binary (CWB) WR140 (HD 193793), observed in 2008, before its periastron passage in 2009, and in 2016–17, spanning the subsequent periastron passage. Both absorption and emission components showed strong variations. The variation of the absorption component as the O5 star was occulted by the wind-collision region (WCR) sets a tight constraint on its geometry. While the sightline to the O5 star traversed the WCR, the strength and breadth of the absorption component varied significantly on time-scales of days. An emission sub-peak was observed on all our profiles. The variation of its radial velocity with orbital phase was shown to be consistent with formation in the WCR as it swung round the stars in their orbit. Modelling the profile gave a measure of the extent of the sub-peak forming region. In the phase range 0.93–0.99, the flux in the sub-peak increased steadily, approximately inversely proportionally to the stellar separation, indicating that the shocked gas in the WCR where the line was formed was adiabatic. After periastron, the sub-peak flux was anomalously strong and varied rapidly, suggesting formation in clumps down-stream in the WCR. For most of the time, its flux exceeded the 2–10-keV X-ray emission, showing it to be a significant coolant of the shocked wind.

Peredur M Williams↗
NASA NTRS2021

Conditions in the WR 140 Wind-Collision Region Revealed by the 1.083-mu m Hei Line Profile

We present spectroscopy of the P Cygni profile of the 1.083-μm He I line in the WC7+O5 colliding-wind binary (CWB) WR 140 (HD 193793), observed in 2008, before its periastron passage in 2009, and in 2016–2017, spanning the subsequent periastron passage. Both absorption and emission components showed strong variations. The variation of the absorption component as the O5 star was occulted by the wind-collision region (WCR) sets a tight constraint on its geometry. While the sightline to the O5 star traversed the WCR, the strength and breadth of the absorption component varied significantly on time-scales of days. An emission subpeak was observed on all our profiles. The variation of its radial velocity with orbital phase was shown to be consistent with formation in the WCR as it swung round the stars in their orbit. Modelling the profile gives a measure of the extent of the subpeak-forming region. In the phase range 0.93–0.99, the flux in the subpeak increased steadily, approximately inversely proportionally to the stellar separation, indicating that the shocked gas in the WCR where the line was formed was adiabatic. After periastron, the subpeak flux was anomalously strong and varied rapidly, suggesting formation in clumps downstream in the WCR. For most of the time, its flux exceeded the 2–10-keV X-ray emission, showing it to be a significant coolant of the shocked wind.

Peredur M Williams↗
NASA NTRSDate not supplied

Hall Magnetohydrodynamic Power Generation and Drag Augmentation Using A Coaxial-Electrode Configuration During Hypersonic Entry

For interplanetary missions, a spacecraft must reduce its relative velocity from orbital speeds of multiple kilometers per second to zero in order to safely land on the surface of a planetary body. During this phase known as planetary reentry, the spacecraft undergoes high speed above Mach 5 up to Mach 30. This hyper-sonic flight regime imposes a strong bow shock in front of the vehicle that imparts extreme aerodynamic and thermal effects. In addition, during portions of this regime, a plasma is formed in the post-shock flow-field around the vehicle due to these high temperatures. Interaction between an applied magnetic field with this electrically conductive fluid, or magnetohydrodynamic (MHD) interaction, can be utilized to convert kinetic energy from the flow into storable electrical energy. Onboard MHD energy generation can benefit the spacecraft by increasing the power of active thermal control components and vehicle control systems during reentry. MHD interaction also exerts a Lorentz body force on the plasma flow that can be manipulated to act as an additional drag force through the vehicle body. MHD drag augmentation can benefit the spacecraft with an additional control mechanism that operates without moving parts. This can be utilized at higher altitudes and subsequently reduce convective heat transfer to the vehicle. MHD interaction with plasma flow suggests that both power generation and drag augmentation can be integrated together in one design. Modern numerical studies have only demonstrated the viability and potential of MHD energy generation and drag modulation separately for implementation into future spacecraft designs [1] [2] [3]. Steeves et. al. presented one possible generator design specifically for reentry vehicles that utilized a modular panel with two embedded electromagnets and two extruding electrodes for air plasma to flow between. Fujino et. al. presented results that indicated increased total drag with a permanent magnet embedded in the blunt body of the Orbital Reentry Experiment (OREX) trajectory. Furthermore, experimental investigations of these two MHD applications have been limited and separate due to the difficult technical nature of experimental de-sign and analysis. For MHD energy generation, there have been two designs which used a cylindrical fore-body with two embedded permanent magnets and two outward facing curved electrodes for tangency to an artificially ionized plasma that flowed along the body length [4] [5]. The first design was studied using an air microwave (MW) ionized supersonic plasma wind tunnel and the second design used radio frequency (RF) ionization of Argon, but both physically demonstrated the feasibility of MHD energy generation in reentry plasma conditions. For MHD drag modulation, one study involved a small-scale cylindrical blunt body embedded with permanent magnets immersed in plasma flow in an arcjet tunnel [6], and another utilized a spherical permanent magnet enclosed with a spherical forebody that was evaluated in a shock expansion tunnel [7]. Both experimental studies provided physical evidence of drag augmentation due to MHD interactions. Therefore, because of the multi-faceted effects of MHD along with both applications having been deemed feasible and beneficial during reentry, this indicates that a coupled power generation and drag modulation design could be achievable. The axi-symmetric nature of blunt bodies for reentry suggests that a spherical set of two ring electrodes near the nose along the body axis. Based on first principles, flow along the walls of the spherical body can then produce MHD interactions between the electrodes. As a result, two currents are generated: a Hall current that produces power and a primary inductive current that augments the drag force. The goal of this work is to provide a simulated design of a coaxial-electrode con-figuration for power generation and drag modulation for experimental testing.

E. Leong↗
NASA NTRS1971

Optimum instantaneous impulsive orbital injection to attain a specified asymptotic velocity vector.

A nalysis of the necessary conditions of Battin for instantaneous orbital injection, with consideration of the uniqueness of his solution, and of the further problem which arises in the degenerate case when radius vector and asymptotic vector are separated by 180 deg. It is shown that when the angular separation between radius vector and asymptotic velocity vector satisfies theta not equal to 180 deg, there are precisely two insertion-velocity vectors which permit attainment of the target asymptotic velocity vector, one yielding posigrade, the other retrograde motion. When theta equals to 180 deg, there is a family of insertion-velocity vectors which permit attainment of a specified asymptotic velocity vector with a unique insertion-velocity vector for every arbitrary orientation of a target unit angular momentum vector.

Bean, W. C.↗
NASA NTRS1980

Mission analysis data for inclined geosynchronous orbits, part 1

Data needed for preliminary design of inclined geosynchronous missions are provided. The inertial and Earth fixed coordinate systems are described, as well as orbit parameters and elements. The complete family of geosynchronous orbits is discussed. It is shown that circular inclined geosynchronous orbits comprise only one set in this family. The major orbit perturbation and their separate effects on the geosynchronous orbit are discussed. Detailed information on the orbit perturbation of inclined circular geosynchronous orbits is given, with emphasis on time history data of certain orbital elements. Orbit maintenance delta velocity (V) requirements to counteract the major orbit perturbations are determined in order to provide order of magnitude estimates and to show the effects of orbit inclination on delta V. Some of the considerations in mission design for a multisatellite system, such as a halo orbit constellation, are discussed.

Graf, O. F., Jr.↗
NASA NTRS1993

On the location of the oscillations in AE Aquarii

There is much recent interest on the nature of the oscillations in the cataclysmic variable AE Aquarii. Detected in the optical, UV, X-ray, and possibly gamma-ray bands, the origin of these 33-sec oscillations remains poorly understood; the oscillation time-delay curve is phase shifted by about 60 deg with respect to the emission-line radial velocity curve, although both should track the orbital motion of the white dwarf. We present simultaneous absorption and emission-line radial velocities, and from the absorption lines we derive an improved ephemeris. Our emission-line velocities are phase shifted by about 75 deg, and hence are unreliable tracers of the orbital motion of the white dwarf. However, the oscillation orbit is shifted by only 5 deg +/- 3 deg percent confirming that the oscillations arise from a region concentric with the white dwarf. Thus the arrival times of the oscillations yield a measure of the white dwarfs apparent orbital velocity, independent of emission-line radial velocity measurements.

Welsh, William F.↗
NASA NTRS1992

Orbital Transfer Vehicle Engine Technology High Velocity Ratio Diffusing Crossover

High speed, high efficiency head rise multistage pumps require continuous passage diffusing crossovers to effectively convey the pumped fluid from the exit of one impeller to the inlet of the next impeller. On Rocketdyne's Orbital Transfer Vehicle (OTV), the MK49-F, a three stage high pressure liquid hydrogen turbopump, utilizes a 6.23 velocity ratio diffusing crossover. This velocity ratio approaches the diffusion limits for stable and efficient flow over the operating conditions required by the OTV system. The design of the high velocity ratio diffusing crossover was based on advanced analytical techniques anchored by previous tests of stationary two-dimensional diffusers with steady flow. To secure the design and the analytical techniques, tests were required with the unsteady whirling characteristics produced by an impeller. A tester was designed and fabricated using a 2.85 times scale model of the MK49-F turbopumps first stage, including the inducer, impeller, and the diffusing crossover. Water and air tests were completed to evaluate the large scale turbulence, non-uniform velocity, and non-steady velocity on the pump and crossover head and efficiency. Suction performance tests from 80 percent to 124 percent of design flow were completed in water to assess these pump characteristics. Pump and diffuser performance from the water and air tests were compared with the actual MK49-F test data in liquid hydrogen.

Lariviere, Brian W.↗
NASA NTRS1987

Study of the mechanism for solar wind formation

Observations of the corona and solar wind are analyzed and compared with generalized results derived from laboratory-scale experiments. It was shown that a thermal pressure gradient can make a major contribution to a precipitating plasma of the solar wind emanating from coronal holes. It is found that the divergence Phi = (R/R sub solar radius)f of the magnetic field lines, originating from coronal holes, is one of the factors governing solar wind velocity at Earth orbit (R= 1 AU). A decrease in the velocity V sub R = 1 AU from approx = 750 mk/sec down to approx = 450 km/sec may be attributable to an increase in superradial divergence f from approx = 7-9 to 20. The plasma energy flux density F at the base of the coronal holes representing the sources of the solar wind with V sub R=1AE = (450 to 750) km/sec, remains nearly constant, being F approx = (1.4 +/- 0.3) x 10 to the 6th power x ergs/sq cm/sec for the period 1973-1975.

Eselevich, V. G.↗
NASA NTRS2016

Design and Stability of an On-Orbit Attitude Control System Using Reaction Control Thrusters

NASA is providing preliminary design and requirements for the Space Launch System Exploration Upper Stage (EUS). The EUS will provide upper stage capability for vehicle ascent as well as on-orbit control capability. Requirements include performance of on-orbit burn to provide Orion vehicle with escape velocity. On-orbit attitude control is accommodated by a on-off Reaction Control System (RCS). Paper provides overview of approaches for design and stability of an attitude control system using a RCS.

Hall, Robert A.↗
NASA NTRS1983

Capture-ejector satellites

A satellite in the form of a large rotating rim which can be used to boost spacecraft from low-Earth orbit to higher orbits is described. The rim rotates in the plane of its orbit such that the lower portion of the rim is traveling at suborbital velocity, while the upper portion is travelling at greater than orbital velocity. Ascending spacecraft or payloads arrive at the lowest portion of the rim at suborbital velocities, where the payloads are released on a trajectory for higher orbits; descending payloads employ the reverse procedure. Electric thrusters placed on the rim maintain rim rotational speed and altitude. From the standpoint of currently known materials, the capture-ejector concept may be useful for relatively small velocity increments.

Macconochie, I. O.↗
NASA NTRS1987

Nuclear powered Mars cargo transport mission utilizing advanced ion propulsion

Nuclear-powered ion propulsion technology was combined with detailed trajectory analysis to determine propulsion system and trajectory options for an unmanned cargo mission to Mars in support of manned Mars missions. A total of 96 mission scenarios were identified by combining two power levels, two propellants, four values of specific impulse per propellant, three starting altitudes, and two starting velocities. Sixty of these scenarios were selected for a detailed trajectory analysis; a complete propulsion system study was then conducted for 20 of these trajectories. Trip times ranged from 344 days for a xenon propulsion system operating at 300 kW total power and starting from lunar orbit with escape velocity, to 770 days for an argon propulsion system operating at 300 kW total power and starting from nuclear start orbit with circular velocity. Trip times for the 3 MW cases studied ranged from 356 to 413 days. Payload masses ranged from 5700 to 12,300 kg for the 300 kW power level, and from 72,200 to 81,500 kg for the 3 MW power level.

Galecki, Diane L.↗
NASA NTRS1987

Nuclear powered Mars cargo transport mission utilizing advanced ion propulsion

Nuclear-powered ion propulsion technology was combined with detailed trajectory analysis to determine propulsion system and trajectory options for an unmanned cargo mission to Mars in support of manned Mars missions. A total of 96 mission scenarios were identified by combining two power levels, two propellants, four values of specific impulse per propellant, three starting altitudes, and two starting velocities. Sixty of these scenarios were selected for a detailed trajectory analysis; a complete propulsion system study was then conducted for 20 of these trajectories. Trip times ranged from 344 days for a xenon propulsion system operating at 300 kW total power and starting from lunar orbit with escape velocity, to 770 days for an argon propulsion system operating at 300 kW total power and starting from nuclear start orbit with circualr velocity. Trip times for the 3 MW cases studied ranged from 356 kW to 413 days. Payload masses ranged from 5700 to 12,300 kg for the 300 kW power level, and from 72,200 to 81, 500 kg for the 3 MW power level.

Galecki, Diane L.↗
NASA NTRS1982

A review of aeroassisted orbit transfer

The development of a theory of aeroassisted orbital plane change maneuvers and necessary equipment is traced and an assessment is made of potential applications. Synergetic plane changes are effected through a combination of aerodynamic and propulsive forces involving the dip of an orbiting vehicle into the atmosphere, performance of an aerodynamic turn using thrust if necessary, then reboosting into a reconfigured orbit. The Orbiter is noted to possess large synergetic plane change capability, and configurations of spacecraft such as the Venus Orbiting Imaging radar for using aeroassist to brake interplanetary velocities before establishing orbits are described. Technology advances necessary to construct effective aerobrake, aerocapture, and orbital transfer vehicles are explored, including fabrication of inflatable ballutes, lighter L/D configurations, and lifting brakes, which are considered to be minor extensions of current capabilities.

Walberg, G. D.↗
NASA NTRS1976

Optimal rendezvous in the neighborhood of a circular orbit

The minimum velocity-change rendezvous solutions, when the motion may be linearized about a circular orbit, fall into two separate regions; the phase-for-free region and the general region. Phase-for-free solutions are derived from the optimum transfer solutions, require the same velocity-change expenditure, but may not be unique. Analytic solutions are presented in two of the three subregions. An algorithm is presented for determining the unique solutions in the general region. Various sources of initial conditions are discussed and three examples are presented.

Jones, J. B.↗
NASA NTRS1982

Capture-ejector satellites

A satellite in the form of a large rotating rim is described which can be used to boost spacecraft from low-Earth orbit to higher orbits. The rim rotates in the plane of its orbit such that the lower portion of the rim is travelling at suborbital velocity, while the upper portion is travelling at greater than orbital velocity. Ascending spacecraft or payloads arrive at the lowest portion of the rim at suborbital velocities, attach to the perimeter, and remain until they reach the highest point, where the payloads are released on a trajectory for higher orbits; descending payloads employ the reverse procedure. Electric thrusters placed on the rim maintain rim rotational speed and altitude. From the standpoint of currently known materials, the capture-ejector concept may be useful for relatively small velocity increments.

Macconochie, I. O.↗
NASA NTRS1987

Sample levitation and melt in microgravity

A system is described for maintaining a sample material in a molten state and away from the walls of a container in a microgravity environment, as in a space vehicle. A plurality of sources of electromagnetic radiation, such as of an infrared wavelength, are spaced about the object, with the total net electromagnetic radiation applied to the object being sufficient to maintain it in a molten state, and with the vector sum of the applied radiation being in a direction to maintain the sample close to a predetermined location away from the walls of a container surrounding the sample. For a processing system in a space vehicle that orbits the Earth, the net radiation vector is opposite the velocity of the orbiting vehicle.

Moynihan, Philip I.↗
NASA NTRS1990

Sample levitation and melt in microgravity

A system is described for maintaining a sample material in a molten state and away from the walls of a container in a microgravity environment, as in a space vehicle. A plurality of sources of electromagnetic radiation, such as an infrared wavelength, are spaced about the object, with the total net electromagnetic radiation applied to the object being sufficient to maintain it in a molten state, and with the vector sum of the applied radiation being in a direction to maintain the sample close to a predetermined location away from the walls of a container surrounding the sample. For a processing system in a space vehicle that orbits the Earth, the net radiation vector is opposite the velocity of the orbiting vehicle.

Moynihan, Philip I.↗