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At least 415 records · Page 23

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NASA NTRS1993

Adjustable Bracket For Entry Of Welding Wire

Wire-entry bracket on welding torch in robotic welding system provides for adjustment of angle of entry of welding wire over range of plus or minus 30 degrees from nominal entry angle. Wire positioned so it does not hide weld joint in view of through-the-torch computer-vision system part of robot-controlling and -monitoring system. Swiveling bracket also used on nonvision torch on which wire-feed-through tube interferes with workpiece. Angle simply changed to one giving sufficient clearance.

Gilbert, Jeffrey L.↗

NASA NTRS1993

Aerothermodynamic performance and thermal protection design for blunt re-entry bodies at L/D = 0.3

Aerodynamic heating and thermal protection design analyses were performed for three blunt re-entry bodies at an L/D = 0.3 returning from low earth orbit. These configurations consisted of a scaled up Apollo command module, a Viking re-entry vehicle, and an Aeroassist Flight Experiment (AFE) aerobrake, each with a maximum diameter of 4.42 m. The aerothermodynamic analysis determined the equilibrium stagnation point heating rate and heat load for nominal and 3-sigma re-entry trajectories and the distribution of heating along the pitch and yaw planes for each of the vehicles at the time of highest heat flux. Using the predicted heating rates and heating distributions, a Thermal Protection System (TPS) design with flight certified materials was tailored for each of the configurations. Results indicated that the heating to the corner of the Viking aeroshell would exceed current limits of reusable tile material. Also, the maximum heating for the AFE would be 15 percent greater than the maximum heating for the Apollo flying the same trajectory. TPS designs showed no significant advantage in TPS weight between the different vehicles; however, heat-shield areal density comparisons showed the Apollo configuration to be the most efficient in terms of TPS weight.

Caram, Jose M.↗

NASA NTRS1993

Atmospheric entry of Mars-return nuclear-powered vehicles due to accidental termination of operations

The entry of nuclear reactors into Earth's atmosphere resulting from an accidental or inadvertent abort of a space vehicle powered by nuclear-thermal rockets is investigated. The study is made for a typical piloted Mars mission vehicle incapacitated by an accident or malfunction during the Earth-arrival phase of the Mars-return journey due to simultaneous, multiple failures of its component systems. A single accident/abort scenario resulting in three entry possibilities is considered for a nominal hyperbolic in-bound approach velocity of 8 km/sec. The most severe case involving a direct entry is then analyzed over a broad range of approach velocities extending to 12 km/sec to include sprint-type missions. The results indicate that the severe surface heating, stagnation pressures, and g-loads are greater than 150 kW/sq cm, 300 atm, and 800-g, respectively. The wall heat transfer rate exceeds the value that can be accommodated by a carbon heatshield through radiation equilibrium prior to sublimation at 5500 K. These conditions are beyond our previous experience in crew safety, structural design, and thermal protection.

Menees, Gene P.↗

NASA NTRS1993

Six-degree-of-freedom guidance and control-entry analysis of the HL-20

The ability of the HL-20 lifting body to fly has been evaluated for an automated entry from atmospheric interface to landing. This evaluation was required to demonstrate that not only successful touchdown conditions would be possible for this low lift-to-drag-ratio vehicle, but also the vehicle would not exceed its design dynamic pressure limit of 400 psf during entry. This dynamic pressure constraint limit, coupled with limited available pitch-control authority at low supersonic speeds, restricts the available maneuvering capability for the HL-20 to acquire the runway. One result of this analysis was that this restrictive maneuvering capability does not allow the use of a model-following atmospheric entry-guidance algorithm, such as that used by the Space Shuttle, but instead requires a more adaptable guidance algorithm. Therefore, for this analysis, a predictor-corrector guidance algorithm was developed that would provide successful touchdown conditions while not violating the dynamic pressure constraint. A flight-control system was designed and incorporated, along with the predictor-corrector guidance algorithm, into a six-DOF simulation. which showed that the HL-20 remained controllable and could reach the landing site and execute a successful landing under all off-nominal conditions simulated.

Powell, Richard W.↗

NASA NTRS1993

Parachute Swivel Mechanism for planetary entry

A parachute swivel mechanism (PSM) for planetary entry missions such as a Mars probe (MARSNET) or return of cometary material samples (ROSETTA mission) has been developed. The purpose of the PSM is to decouple the spin of the probe from the parachute, with low friction torque, during both the deployment and descent phases. Critical requirements are high shock loads, low friction, low temperatures, and several years of storage in the deep space environment (during the cruise phase of the probe, prior to operation). The design uses a main thrust ball bearing to cope with the load requirement and a smaller thrust ball bearing for guiding of the shaft. Except for use on the Viking and Galileo swivels, it appears that this type of bearing has very rarely been employed in space mechanisms, so that little is known of its friction behavior with dry lubrication. A slip ring assembly allows the transfer of electrical power for post-reefing of the parachute. A test program has been conducted covering the environmental conditions of Mars entry and Earth reentry. This paper describes requirement constraints, model missions of planetary entries, a bearing trade-off, analyses performed, design details, the lubrication system, and test results (friction torque versus load/spin rate). In addition, the design of the test rig is addressed.

Birner, R.↗

NASA NTRS1993

Ambulatory blood pressure and heart rate during shuttle flight, entry and landing

Ambulatory blood pressures (BP) and heart rates (HR) were recorded on a series of early Shuttle flights during preflight and pre-entry, entry, landing and egress. There were no significant differences between flight and preflight values during routine activity. Systolic blood pressure was slightly elevated in the deorbit period and systolic and diastolic blood pressure and heart rates were all elevated with onset of gravitoinertial loads and remained so through egress. Two of seven subjects had orthostatic problems in egress but their data did not show significant differences from others except in heart rate. Comparison of this data to that from recent studies show even larger increase in HR/BP values during current deorbit and entry phases which is consistent with increased heat and weight loads imposed by added survival gear. Both value and limitations of ambulatory heart rate/blood pressure data in this situation are demonstrated.

Thornton, W.↗

NASA NTRS1993

Simulated rarefied entry of the Galileo probe into the atmosphere of Jupiter

Flow properties and aerodynamics are computed with a direct simulation Monte Carlo (DSMC) method for rarefied entry of the Galileo Probe into the atmosphere of Jupiter. Accurate predictions of vehicle drag coefficients are needed in order to assess atmospheric properties from the onboard Atmosphere Structure Experiment where highly-sensitive accelerometers will measure the drag force to within 10-6 barr during the initial entry phase at high altitudes. The corresponding flow rarefraction extends from the free molecule limit to the near continuum transition regime (Re less than 1000). Simulation results indicate that C(sub D) varies from 2.1 at the free molecule limit down to 1.6 at Re(infinity) = 1,000. Temperatures, densities, and internal energies throughout the flow field were also computed at each altitude ranging from 735 km to 353 km above the 1 barr level in the Jovian atmosphere. Surface heating and temperatures of the probe were computed directly in the DSMC code by assuming radiative equilibrium. Material response was re-asssessed accurately during entry by accounting for conductivity, heat capacity, and pyrolysis which led to surface material mass efflux several times that of the freestream mass influx. The simulation also accounted for the quantum nature of the rotational energy mode of the dominant atmospheric species H2 through partial internal excitation in the freestream gas.

Hass, Brian L.↗

NASA NTRS1993

Wave-like perturbations observed at low altitudes by the Pioneer Venus Orbiter Neutral Mass Spectrometer during orbiter entry

Wave-like perturbations have been observed in the nightside neutral density data acquired for He, N, O, N2 and CO2 by the Pioneer Venus Orbiter Neutral Mass Spectrometer (ONMS) during entry in late 1992. The data cover an altitude range of 133-200 km from 0.5-4.5 hours local solar time and occur at medium solar activity (F(sub 10.7) = 120). The perturbations, with an effective wavelength along the orbit of about 100 to 600 km, have similar amplitudes for the various species and helium is out of phase with respect to the heavier mass species. The measurements are comparable to those observed in 1978-80 at solar maximum activity (F(sub 10.7) = 200) above 145 km. Between 133 and 160 km the rms amplitudes grow with altitude at a rate of about (1.6, 2.1, 2.7, 4.4) x 10(exp 3)/km for N, O, N2, and CO2, respectively. The average rms amplitudes above 145 km of 0.08 for N and O and 0.1 for N2 are comparable in magnitude to those observed in the earlier 1978-80 data of 0.06, 0.08, and 0.095 respectively. CO2 is an exception for which the entry value is 0.17 compared to 0.09 earlier. By combining the two overlapping data sets there is a suggestion that the CO2 amplitudes grow in value with altitude up to about 140-170 km and then decrease in amplitude. Like the earlier data, the entry data are consistent with the interpretation that the neutral density perturbations are due to gravity waves propagating upward from the lower thermosphere.

Kasprzak, W. T.↗

NASA NTRS1981

Entry and Landing Probe for Titan

Results of a recent study of entry and landing probes for the exploration of Titan are presented. The probes considered were based on a wide range of exploration mission possibilities. They included: an atmospheric science probe, an intermediate atmospheric and limited surface science probe, and a larger atmospheric and expanded surface science probe. Because of lower gravity on Titan and its atmospheric characteristics, the entry environment is less severe than that of Mars. However, the large uncertainties in the current definition of the atmosphere and the uncertainties in Titan's surface characteristics have required tradeoffs of various combinations of entry and descent shapes and hard lander configurations. Results show that all probe classes are feasible without major developments.

Murphy, J. P.↗

NASA NTRS1960

Investigation of the Errors of an Inertial Guidance System during Satellite Re-Entry

During the re-entry phase of a manned satellite, some equipment for continuous on-board indication of position will be required. Since radio and radar may be useless during part of the re-entry, inertial guidance equipment may be required. Such equipment, however, has an inherent instability in the computation of altitude. The present study of an inertial guidance system shows that for reasonable values of initial-condition errors and equipment biases. the resultant position indication errors will not become excessive unless the re-entry maneuver time is greater than 45 minutes to an hour. Further, the position indication error caused by accelerometer uncertainties can be reduced by switching off the accelerometers until their output becomes significantly greater than their uncertainty.

White, John S.↗

NASA NTRS1961

The Use of Drag Modulation to Limit the Rate at Which Deceleration Increases During Nonlifting Entry

The method developed in NASA TN D-319 for studying the atmosphere entry of vehicles with varying aerodynamic forces has been applied to obtain a closed-form solution for the motion, heating, range, and variation of the vehicle parameter m/C(D)A for nonlifting entries during which the rate of increase of deceleration is limited. The solution is applicable to vehicles of arbitrary weight, size, and shape, and to arbitrary atmospheres. Results have been obtained for entries into the earth's atmosphere at escape velocity during which the maximum deceleration and the rate at which deceleration increases were limited. A comparison of these results with those of NASA TN D-319, in which only the maximum deceleration was limited, indicates that for a given corridor depth, limiting the rate of increase of deceleration and the maximum deceleration requires an increase in the magnitude of the change in M/C(D)A and results in increases in maximum heating rate, total heat absorbed at the stagnation point, and range.

Levy, Lionel L., Jr.↗

NASA NTRS1960

Modulated Entry

The technique of modulation, or variable coefficients, is discussed and the analytical formulation is reviewed. Representative numerical results of the use of modulation are shown for the lifting and nonlifting cases. These results include the effects of modulation on peak acceleration, entry corridor, and heat absorption. Results are given for entry at satellite speed and escape speed. The indications are that coefficient modulation on a vehicle with good lifting capability offers the possibility of sizable loading reductions or, alternatively, wider corridors; thus, steep entries become practical from the loading standpoint. The amount of steepness depends on the acceptable heating penalty. The price of sizable fractions of the possible gains does not appear to be excessive.

Grant, Frederick C.↗

NASA NTRS1993

Geophysical Research Letters. Selected Papers on Pioneer Venus Orbiter: Entry Phase

Contents include the following papers which are comprised of subject matter related to the The Pioneer Venus Orbiter's Entry Phase: The Pioneer Venus entry phase; solar cycle variations of electron density and temperature in the Venusian nightside ionosphere; the magnetic state of the lower ionosphere during Pioneer Venus entry phase; the nightside ionosphere of Venus under varying levels of solar EUV flux; observation of the nightside Venus ionosphere; final encounter of the Pioneer Venus orbiter ion mass spectrometer; ion measurement during Pioneer Venus reentry; implications for solar cycle variation of ion composition and dynamics; evidence for day-to-night ion transport at low solar activity in the Venus pre-dawn ionosphere; model calculations of the dayside ionosphere of Venus at solar minimum; natural composition measurements by the Pioneer Venus neutral mass spectrometer during orbiter reentry; the Venus atmospheric response to solar cycle variations; and plasma waves observed at low altitudes in the tenuous Venus nightside ionosphere.

Source record↗

NASA NTRS1997

Guidance and Control for Mars Atmospheric Entry: Adaptivity and Robustness

In this paper, we address the atmospheric entry guidance and control problem for Mars precision landing. The guidance and control design is based on the principle of tracking a reference drag versus velocity profile in the entry flight corridor, which is determined by physical constraints of the flight. An integrated adaptive/robust control approach to atmospheric entry guidance and control is introduced to deal with different uncertainties.

Lu, Wei-Min↗

NASA NTRS2001

An Earth Entry Vehicle For Returning Samples From Mars

The driving requirement for design of a Mars Sample return mission is assuring containment of the returned samples. The impact of this requirement on developmental costs, mass allocation, and design approach of the Earth Entry Vehicle is significant. A simple Earth entry vehicle is described which can meet these requirements and safely transport the Mars Sample Return mission's sample through the Earth's atmosphere to a recoverable location on the surface. Detailed analysis and test are combined with probabilistic risk assessment to design this entirely passive concept that circumvents the potential failure modes of a parachute terminal descent system. The design also possesses features that mitigate other risks during the entry, descent, landing and recovery phases. The results of a full-scale drop test are summarized.

Mitcheltree, R.↗

NASA NTRS2000

Earth Entry Vehicle for Mars Sample Return

Mitcheltree, R. A.↗

NASA NTRS2001

Integrated Thermal Response Tool for Earth Entry Vehicles

A system is presented for multi-dimensional, fully-coupled thermal response modeling of hypersonic entry vehicles. The system consists of a two-dimensional implicit thermal response, pyrolysis and ablation program (TITAN), a commercial finite-element thermal and mechanical analysis code (MARC), and a high fidelity Navier-Stokes equation solver (GIANTS). The simulations performed by this integrated system include hypersonic flow-field, fluid and solid interaction, ablation, shape change, pyrolysis gas generation and flow, and thermal response of heatshield and structure. The thermal response of the ablating and charring heatshield material is simulated using TITAN, and that of the underlying structural is simulated using MARC. The ablating heatshield is treated as an outer boundary condition of the structure, and continuity conditions of temperature and heat flux are imposed at the interface between TITAN and MARC. Aerothermal environments with fluid and solid interaction are predicted by coupling TITAN and GIANTS through surface energy balance equations. With this integrated system, the aerothermal environments for an entry vehicle and the thermal response of both the heatshield and the structure can be obtained simultaneously. Representative computations for a proposed blunt body earth entry vehicle are presented and discussed in detail.

Chen, Y.-K.↗

NASA NTRS2003

Multibody Parachute Flight Simulations for Planetary Entry Trajectories Using "Equilibrium Points"

A method has been developed to reduce numerical stiffness and computer CPU requirements of high fidelity multibody flight simulations involving parachutes for planetary entry trajectories. Typical parachute entry configurations consist of entry bodies suspended from a parachute, connected by flexible lines. To accurately calculate line forces and moments, the simulations need to keep track of the point where the flexible lines meet (confluence point). In previous multibody parachute flight simulations, the confluence point has been modeled as a point mass. Using a point mass for the confluence point tends to make the simulation numerically stiff, because its mass is typically much less that than the main rigid body masses. One solution for stiff differential equations is to use a very small integration time step. However, this results in large computer CPU requirements. In the method described in the paper, the need for using a mass as the confluence point has been eliminated. Instead, the confluence point is modeled using an "equilibrium point". This point is calculated at every integration step as the point at which sum of all line forces is zero (static equilibrium). The use of this "equilibrium point" has the advantage of both reducing the numerical stiffness of the simulations, and eliminating the dynamical equations associated with vibration of a lumped mass on a high-tension string.

Raiszadeh, Ben↗