SEARCH · Search NASA

Results for “Orbiter Loads”

Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

SourceResearch connectionFrom yearThrough year

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 55 records · Page 3

All sources NASA NTRS DOE OSTI

NASA NTRS1976

Filament wound pressure vessels with load sharing liners for Space Shuttle Orbiter applications

It is recognized that the use of overwrapped pressure vessels with load sharing liners may provide significant weight savings for high pressure gas containment in Space Shuttle Orbiter systems. The technology readiness to produce Kevlar wound vessels with load sharing liners of titanium 6Al-4V, Inconel 718 or cryoformed 301 steel has been demonstrated. It has been estimated that about 400 lbs can be saved in the Orbiter by using overwrapped vessels with load sharing liners instead of monolithic metal designs. Total weight of the composite vessels would be about 1350 lbs as opposed to about 1750 lbs for all-metal vessels.

Ecord, G. M.↗

NASA NTRS2019

The Exo-Brake as an Inexpensive Means of Achieving Sample Return from Low Earth Orbit – Recent Flight

The Exo-Brake is a simple, non-propulsive means of deorbiting small pay-loads from orbital platforms such as the Inter-national Space Station (ISS). Recent flight experiments involving the TechEdSat (TES) 6, 7, 8 are discussed in terms of both ‘targeted’ and ‘disposal’ de-orbit techniques. These build on the previous flight experiments with fixed surface areas – and now involve improved uplink/downlink communication and GPS for improved targeting and control. The recent targeting experiments are dis-cussed involving the TechEdSat-6,7,8 nanosatellites. The extension of the concept to a 1-stage, 3-stage, and lifting entry sample re-turn system are discussed.

Murbach, Marcus S.↗

NASA NTRS2008

Simulated Reentry Heating by Torching

The two first order reentry heating parameters are peak heating flux (W/cm2) and peak heat load (kJ/cm2). Peak heating flux (and deceleration, gs) is higher for a ballistic reentry and peak heat load is higher for a lifting reentry. Manned vehicle reentries are generally lifting reentries at nominal 1-5 gs so that personnel will not be crushed by high deceleration force. A few off-nominal manned reentries have experienced 8 or more gs with corresponding high heating flux (but below nominal heat load). The Shuttle Orbiter reentries provide about an order of magnitude difference in peak heating flux at mid-bottom (TPS tiles, approximately 6 W/cm2 or 5 BTU/ft2- sec) and leading edge (RCC, approximately 60 W/cm2 or 50 BTU/ft2- sec). Orion lunar return and Mars sample lander are of the same order of magnitude as orbiter leading edge peak heat loads. Flight temperature measurements are available for some orbiter TPS tile and RCC locations. Return-to-Flight on-orbit tile-repair-candidate-material-heating performance was evaluated by matching propane torch heating of candidate-materials temperatures at several depths to orbiter TPS tile flight-temperatures. Char and ash characteristics, heat expansion, and temperature histories at several depths of the cure-in-place ablator were some of the TPS repair material performance characteristics measured. The final char surface was above the initial surface for the primary candidate (silicone based) material, in contrast to a receded surface for the Apollo-type ablative heat shield material. Candidate TPS materials for Orion CEV (LEO and lunar return), and for Mars sample lander are now being evaluated. Torching of a candidate ablator material, PICA, was performed to match the ablation experienced by the STARDUST PICA heat shield. Torching showed that the carbon fiberform skeleton in a sample of PICA was inhomogeneous in that sample, and allowed measurements (of the clumps and voids) of the inhomogeneity. Additional reentry heating-performance characterizations of high temperature insulation materials were performed.

Harvey, Gale A.↗

NASA NTRS2008

Simulated Reentry Heating by Torching

The two first order reentry heating parameters are peak heating flux (W/square cm) and peak heat load (kJ/square cm). Peak heating flux (and deceleration, gs) is higher for a ballistic reentry and peak heat load is higher for a lifting reentry. Manned vehicle reentries are generally lifting reentries at nominal 1-5 gs so that personnel will not be crushed by high deceleration force. A few off-nominal manned reentries have experienced 8 or more gs with corresponding high heating flux (but below nominal heat load). The Shuttle Orbiter reentries provide about an order of magnitude difference in peak heating flux at mid-bottom (TPS tiles, approximately 6 W/square cm or 5 BTU/square ft - sec) and leading edge (RCC, approximately 60 W/square cm or 50 BTU/square ft- sec). Orion lunar return and Mars sample lander are of the same order of magnitude as orbiter leading edge peak heat loads. Flight temperature measurements are available for some orbiter TPS tile and RCC locations. Return-to-Flight on-orbit tile-repair-candidate-material-heating performance was evaluated by matching propane torch heating of candidate-materials temperatures at several depths to orbiter TPS tile flight-temperatures. Char and ash characteristics, heat expansion, and temperature histories at several depths of the cure-in-place ablator were some of the TPS repair material performance characteristics measured. The final char surface was above the initial surface for the primary candidate (silicone based) material, in contrast to a receded surface for the Apollo-type ablative heat shield material. Candidate TPS materials for Orion CEV (LEO and lunar return), and for Mars sample lander (MSL) are now being evaluated. Torching of a candidate ablator material, PICA, was performed to match the ablation experienced by the STARDUST PICA heat shield. Torching showed that the carbon fiberform skeleton in a sample of PICA was inhomogeneous in that sample, and allowed measurements (of the clumps and voids) of the inhomogeneity. Additional reentry heating-performance characterizations of high temperature insulation materials were performed.

Harvey, Gale A.↗

NASA NTRS1977

Static aerodynamic characteristics of a single-stage-to-orbit vehicle with low planform loading at Mach numbers from 0.3 to 4.63

Transonic pressure and wind tunnel studies were performed to determine the longitudinal and lateral aerodynamic characteristics of a single-stage-to-orbit vehicle which utilizes an all metallic, hot structure, thermal protection system resulting in low planform loading. The model was tested over a Mach number range from 0.3 to 4.63 for angles of attack from -4 deg to 32 deg at both 0 deg and 5 deg sideslip.

Freeman, D. C., Jr.↗

NASA NTRS1974

Electrical power generation subsystem for Space Shuttle Orbiter

The requirements, capabilities, and design of the Space Shuttle Orbiter electrical power generation (EPG) subsystem are presented. The Orbiter EPG subsystem is designed to have the flexibility to supply the basic Orbiter electrical loads and the power requirements of the payloads which, in some cases, are completely dependent on the Orbiter for both power and heat rejection. These needs are supplied by three hydrogen/oxygen fuel cell powerplants (FCPs), having the capability of providing a total of 14 kW average and up to 24 kW peak in the basic Orbiter configuration. Kits permit dedication of one FCP to the payload by providing an additional reactant tank pair for a seven-day mission and additional heat rejection capability.

Blaski, M. F.↗

NASA NTRS1974

Effect of gaseous and solid simulated jet plumes on an 040A space shuttle launch configuration at m=1.6 to 2.2

The effect of plume-induced flow separation and aspiration effects due to operation of both orbiter and the solid rocket motors on a 0.019-scale model of the launch configuration of the Space Shuttle Vehicle is determined. Longitudinal and lateral-directional stability data were obtained at Mach numbers of 1.6, 2.0, and 2.2 with and without the engines operating. The plumes exiting from the engines were simulated by a cold-gas jet supplied by an auxiliary 200-atm air supply system and solid-body plume simulators. The aerodynamic effects produced by these two simulation procedures are compared. The parameters most significantly affected by the jet plumes are pitching moment, elevon control effectiveness, axial force, and orbiter wing loads. The solid rocket motor (SRM) plumes have the largest effect on the aerodynamic characteristics. The effect of the orbiter plumes in combination with the SRM plumes is also significant. Variations in the nozzle design parameters and configuration changes can reduce the jet plume-induced aerodynamic effects.

Dods, J. B., Jr.↗

NASA NTRS1988

Low-Heat-Transfer Tank Mount

Supporting element adapts to high or low side loads. Passive Orbital Disconnect Strut (PODS) for cryogenic tank engages an extra-strong member only during peak side or axial loads like those from vibrations, accelerations, and shocks. When side loads are low, support provided by lower-strength, lower-thermal-conductance, passive disconnect.

Parmley, R. T.↗

NASA NTRS1998

An Adaptable Power System with Software Control Algorithm

A low cost, flexible and modular spacecraft power system design was developed in response to a call for an architecture that could accommodate multiple missions in the small to medium load range. Three upcoming satellites will use this design, with one launch date in 1999 and two in the year 2000. The design consists of modular hardware that can be scaled up or down, without additional cost, to suit missions in the 200 to 600 Watt orbital average load range. The design will be applied to satellite orbits that are circular, polar elliptical and a libration point orbit. Mission unique adaptations are accomplished in software and firmware. In designing this advanced, adaptable power system, the major goals were reduction in weight volume and cost. This power system design represents reductions in weight of 78 percent, volume of 86 percent and cost of 65 percent from previous comparable systems. The efforts to miniaturize the electronics without sacrificing performance has created streamlined power electronics with control functions residing in the system microprocessor. The power system design can handle any battery size up to 50 Amp-hour and any battery technology. The three current implementations will use both nickel cadmium and nickel hydrogen batteries ranging in size from 21 to 50 Amp-hours. Multiple batteries can be used by adding another battery module. Any solar cell technology can be used and various array layouts can be incorporated with no change in Power System Electronics (PSE) hardware. Other features of the design are the standardized interfaces between cards and subsystems and immunity to radiation effects up to 30 krad Total Ionizing Dose (TID) and 35 Mev/cm(exp 2)-kg for Single Event Effects (SEE). The control algorithm for the power system resides in a radiation-hardened microprocessor. A table driven software design allows for flexibility in mission specific requirements. By storing critical power system constants in memory, modifying the system code for other programs is simple. These constants can be altered also by ground command, or in response to an anomolous event. All critical power system functions have backup hardware functions to prevent a software or computer glitch from propagating. A number of battery charge control schemes can be implemented by selecting the proper control terms in the code. The architecture allows the design engineer to tune the system response to various system components and anticipated load profiles without costly alterations. A design trade was made with the size, weight and power dissipation of the electronics versus the performance of the power bus to load variations. Linear, fine control is maintained with a streamlined electronics design. This paper describes the hardware design as well as the software control algorithm. The challenges of closing the system control loop digitally is discussed. Control loop margin and power system performance is presented. Lab measurements are shown and compared to the system response of a hardware model running actual flight software.

Castell, Karen↗

NASA NTRS1992

Space Station Freedom main truss thermal analysis

The main truss structure of the Permanently Manned Capability configuration of the Space Station Freedom is analyzed in order to evaluate orbital heating loads and determine the temperature distributions and temperature gradients in the structure. The results are to be utilized in a thermal/structure analysis in the determination of thermal stresses and deflections in the truss members caused by solar heating loads and by shadowing of the truss members by other segments of the Space Station as it moves around the orbit. A SINDA thermal finite different model and a TRASYS radiation interchange model of the truss were generated automatically from a NASTRAN finite element structural model. The analysis yielded transient temperature distribution in the truss structure around the orbit, indicating large temperature gradients in the structure.

Warren, Andrew H.↗

NASA NTRS1987

New non-geosynchronous orbits for communications satellites to off-load daily peaks in geostationary traffic

The potential for satellites in two orbits, the sun-synchronous 12-hour equatorial orbit (STET) and the apogee at constant time-of-day equatorial orbit (ACE), to off-load peaks in the CONUS geostationary communications traffic is discussed. These orbits are found to require maneuvers of smaller magnitudes for insertion than geostationary orbits. Advantages of the ACE orbit over the STET orbit are discussed, including larger satellite mass capability for a given launch vehicle, lower slant ranges, and larger angular separation from the geostationary arc for a nonequatorial ground observer.

Turner, A. E.↗

NASA NTRS1982

Integrated thermal-structural analysis of large space structures

Optimum performance of large space antennas requires very fine control of the shape of the antenna surface since the shape affects both frequency control and pointing accuracy. A significant factor affecting the antenna shape is the temperature of the structure and the resulting deformation. To accurately predict the temperature of the structure, it is necessary first to accurately predict thermal loads. As the structure orbits the Earth, the thermal loads change constantly so that the thermal-structural response varies continuously throughout the orbit. The results from recent applications of integrated finite element methodology to heat load determination and thermal-structural analysis of large space structures are given. Four areas are concentrated on: (1) the characteristics of the integrated finite element methodology, (2) fundamentals of orbital heat load calculation, (3) description and comparison of some radiation finite elements, and (4) application of the integrated finite-element approach to the thermal-structural analysis of an orbiting truss structure.

Mahaney, J.↗

NASA NTRS1978

Shuttle orbital flight test

The initial orbital flight of the shuttle vehicle will involve several operational aspects which differ from previous manned space programs. The asymmetrical shuttle launch configuration undergoes unique aero and thrust loads. The orbiter requires a complex interactive control system during entry. Techniques are described which will be used to minimize the ascent loads. Procedures are outlined which should maximize the entry control margins. The primary objective of the first shuttle flight is to obtain data from which to extract the loads and stability derivatives. Accomplishment of this objective will enable relaxation of constraints on launch winds, entry center of gravity, and crossrange during subsequent flights.

North, W. J.↗

NASA NTRS2000

Research into the effects of astronaut motion on the spacecraft: a review

The paper reviews the research that has been undertaken to understand and quantify the disturbance effects of the astronaut's motion inside and outside the spacecraft on the vehicle's attitude and acceleratory environment. In early investigations, the dynamic interaction of astronauts, modeled as point masses, and the spacecraft, modelled as a rigid body, was analyzed. Through ground-based experiments and the modeling of astronaut-induced forces and moments as stochastic processes, it became possible to estimate the magnitude and energy content of the loads produced by the astronaut. The first experiment in space to measure the astronaut-induced disturbances was conducted on the Skylab orbital station. Loads generated while performing routine operations were measured on board the Space Shuttle in 1994 and on the space station Mir in 1996-1997. c 2001 Elsevier Science Ltd. All rights reserved.

Skylab Project↗

NASA NTRS1983

Supersonic loads due to shuttle-orbiter/external-tank attachment structures

A large variety of tests were conducted during development of the Space Shuttle to determine the strength, fatigue, and thermal characteristics of the thermal-protection system of the orbiter. The research and analysis carried out to accurately define and simulate the flow characteristics immediately ahead of the external tank attachment structures are described. The highest known acoustic pressure loads occur in these regions and tests to insure the ability of the titles to withstand these acoustic loads, plus those due to pressure gradient, were thought necessary. First flight-critical wind tunnel tests were therefore conducted in the Langley 8-Foot Transonic Pressure Tunnel (8-Ft. TPT), which simulated the time histories of Shuttle ascent loads on tile arrays bonded to structures which accurately duplicated those of the Shuttle. The time-varying free stream conditions were provided by controlling the deflection angle history of diffuser spoiler flaps in an automated way. Time histories of the critical load parameters imposed on the tile arrays in the tunnel are compared with those expected in flight.

Edwards, C. L. W.↗

NASA NTRS2005

Space Shuttle Partial Stack Rollout Test Analytical Correlation In Support Of Fatigue Load Development

A rollout test with only the Solid Rocket Boosters was conducted in November 2003 to gather structural dynamic response data of the transportation environment from the Vehicle Assembly Building to the Launch Pad. The data was acquired to develop and validate analytical methods used to predict rollout Orbiter fatigue load spectra. Earlier predictions computed by a base drive approach with only 5 input drive degrees-of-freedom raised questions that commissioned the partial stack test. Not only was there a concern because of the input degree-of-freedom omission due to measurement limitations, but there was also a concern with the implementation of the "large mass" itself. Three methods were evaluated with the partial stack test data. The analytical correlations to measured strain derived SRB base loads and acceleration5 showed the earlier 5 degree-of-freedom base drive approach to yield the most conservative results for all quantities monitored except the SRB base moment about the axis in which the input drive was missing. This non-conservative shortcoming led to a recommendation to use either the 6 degree-of-freedom base drive or the 12 degree-of-freedom Craig-Bampton boundary drive methods whose results did not substantially differ.

DelBasso, Steve↗

NASA NTRS1973

Results of tests in the MSFC 14 x 14 inch trisonic wind tunnel on a .004 scale model of the Rockwell International Space Shuttle Vehicle 3, (integrated configuration)

Experimental aerodynamic investigations were conducted during mid-July, 1973 on a .004 scale model of the Rockwell International integrated configuration Space Shuttle Vehicle 3. The purpose of the tests was three fold: (1) to determine the static stability characteristics of the integrated vehicle, utilizing the Vehicle 3 orbiter configuration; (2) to determine the effect of interstage structure and tank external fuel lines on the integrated vehicle aerodynamic characteristics; (3) to determine the effects of the aft interstage structure on orbiter aerodynamic loads. Data were recorded on the integrated vehicle (test no. 579) at angles of attack and sideslip ranging from -10 deg to 10 deg over a Mach number schedule from 0.6 to 4.96. Data were obtained on the orbiter alone in the presence of the external tank with SRB attached (test no. 580) at angles of attack from -10 deg to 10 deg over a Mach number range from .6 to 1.96. Plotted data are presented in the body axis system.

Allen, E. C.↗

NASA NTRS1997

Three-Wave Gas Journal Bearing Behavior With Shaft Runout

Experimental orbits of a free-mounted, three-wave gas journal bearing housing were recorded and compared to transient predicted orbits. The shaft was mounted eccentric with a fixed runout. Experimental observations for both the absolute bearing housing center orbits and the relative bearing housing center to shaft center orbits are in good agreement with the predictions. The sub-synchronous whirl motion generated by the fluid film was found experimentally and predicted theoretically for certain speeds. A three-wave journal bearing can run stably under dynamic loads with orbits well inside the bearing clearance. Moreover, the orbits are almost circular free of the influence of bearing wave shape.

Dimofte, Florin↗