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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.

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At least 775 records · Page 43
All sourcesNASA NTRSDOE OSTI
NASA NTRS1990

Optimal autorotational descent of a helicopter with control and state inequality constraints

A point-mass model of the OH-58A helicopter has been used to ascertain the autorotation profiles which minimize helicopter impact velocity while remaining within the bounds of the main rotor's collective pitch and angular speed. The optimal control strategies are comparable to those employed by pilots in autorotational landings. It is noted that a possibility exists for the reduction of the height-sink rate restriction zone of OH-58A helicopters, using optimal energy-management techniques.

Lee, Allan Y.↗
NASA NTRS1990

The 4-D descent trajectory generation techniques under realistic operating conditions

NASA-Langley has been conducting and sponsoring research in airborne energy management for a number of years. During the course of this research, two fundamental techniques for the generation of 4D (fixed time) descent trajectories have emerged as viable candidates for advanced flight management systems. The first technique utilizes speed schedules of constant Mach number transitioning to constant calibrated airspeed chosen empirically to produce minimum fuel usage. The second technique computes cost optimized speed schedules of variable airspeed developed through application of optimal control theory. Both techniques have been found to produce reasonable and flyable descent trajectories. The formulation of the algorithms for each technique is evaluated and their suitability for operations in realistic conditions is discussed. Operational factors considered include: airplace speed, thrust, and altitude rate constaints; wind, temperature, and pressure variations; Air Traffic Control altitude, speed, and time constaints; and pilot interface and guidance considerations. Time flexibility, fuel usage, and airborne computational requirements were the primary performance measures.

Williams, David H.↗
NASA NTRS1991

Personnel launch system autoland development study

The Personnel Launch System (PLS) Autoland Development Study focused on development of the guidance and control system for the approach and landing (A/L) phase and the terminal area energy management (TAEM) phase. In the A/L phase, a straight-in trajectory profile was developed with an initial high glide slope, a pull-up and flare to lower glide slope, and the final flare touchdown. The TAEM system consisted of using a heading alignment cone spiral profile. The PLS autopilot was developed using integral LQG design techniques. The guidance and control design was verified using a nonlinear 6 DOF simulation. Simulation results demonstrated accurate steering during the TAEM phase and adequate autoland performance in the presence of wind turbulence and wind shear.

Bossi, J. A.↗
NASA NTRS1991

X-15 Contributions to the X-30

Some of the less publicized flight test results from the X-15 program that might relate to sustained high-speed flight in the atmosphere are presented. The topics covered include: (1) energy management and range considerations; (2) the advantages of pilot-in-the-loop and redundant-emergency systems; (3) a summary of some of the aerodynamic heating problems that were encountered; and (4) some comments on the advantages of an early flight test program and gradual expansion of the flight envelope.

Hoey, Robert G.↗
NASA NTRS1990

Solid-liquid staged combustion space boosters

NASA has begun to evaluate solid-liquid hybrid propulsion for launch vehicle booster. A three-phase program was outlined to identify, acquire, and demonstrate technology needed to approximate solid and liquid propulsion state of the art. Aerojet has completed a Phase 1 study and recommends a solid-liquid staged combustion concept in which turbopump fed LO2 is burned with fuel-rich solid propellant effluent in aft-mounted thrust chambers.These reasonably sized thrust chambers are LO2 regeneratively cooled, supplemented with fuel-rich barrier cooling. Turbopumps are driven by the resulting GO2 coolant in an expander-bleed-burnoff cycle. Turbine exhaust pressurizes the LO2 tankage directly, and the excess is bled into supersonic nozzle splitlines, where it combusts with the fuel rich boundary layer. Thrust vector control is enhanced by supersonic nozzle movement on flexseal mounts. Every hybrid solid-liquid concept examined improves booster energy management and launch propellant safety compared to current solid boosters. Solid-liquid staged combustion improves hybrid performance by improving both combustion efficiency and combustion stability, especially important for large boosters. These improvements result from careful fluid management and use of smaller combustors. The study shows NASA safety, reliability, cost, and performance criteria are best met with this concept, wherein simple hardware relies on several separate emerging technologies, all of which have been demonstrated successfully.

Culver, D. W.↗
NASA NTRS1989

NASA Space Engineering Research Center for Utilization of Local Planetary Resources

Progress toward the goal of exploiting extraterrestrial resources for space missions is documented. Some areas of research included are as follows: Propellant and propulsion optimization; Automation of propellant processing with quantitative simulation; Ore reduction through chlorination and free radical production; Characterization of lunar ilmenite and its simulants; Carbothermal reduction of ilmenite with special reference to microgravity chemical reactor design; Gaseous carbonyl extraction and purification of ferrous metals; Overall energy management; and Information management for space processing.

Ramohalli, Kumar↗
NASA NTRS1992

STS-46 Space Shuttle mission report

The STS-46 Space Shuttle Program Mission Report contains a summary of the Orbiter, External Tank (ET), Solid Rocket Booster/Redesigned Solid Rocket Motor (SRB/RSRM), and the Space Shuttle main engine (SSME) subsystem performance during the forty-ninth flight of the Space Shuttle Program, and the twelfth flight of the Orbiter vehicle Atlantis (OV-104). In addition to the Atlantis vehicle, the flight vehicle consisted of the following: an ET, designated ET-48 (LWT-41); three SSME's, which were serial numbers 2032, 2033, and 2027 in positions 1, 2, and 3, respectively; and two SRB's which were designated BI-052. The lightweight/redesigned SRM's that were installed in each SRB were designated 360W025A for the left RSRM and 360L025B for the right RSRM. The primary objective of this flight was to successfully deploy the European Retrievable Carrier (EURECA) payload and perform the operations of the Tethered Satellite System-1 (TSS-1) and the Evaluation of Oxygen Interaction with Material 3/Thermal Energy Management Processes 2A-3 (EOIM-3/TEMP 2A-3). The secondary objectives of this flight were to perform the operations of the IMAX Cargo Bay Camera (ICBC), Consortium for Material Development in Space Complex Autonomous Payload-2 and 3 (CONCAP-2 and CONCAP-3), Limited Duration Space Environment Candidate Materials Exposure (LDCE), Pituitary Growth Hormone Cell Function (PHCF), and Ultraviolet Plume Instrumentation (UVPI). In addition to summarizing subsystem performance, this report also discusses each Orbiter, ET, SSME, SRB, and RSRM in-flight anomaly in the applicable section of the report. Also included in the discussion is a reference to the assigned tracking number as published on the Problem Tracking List. All times are given in Greenwich mean time (G.m.t.) as well as mission elapsed time (MET).

Fricke, Robert W.↗
NASA NTRS1992

Proceedings of the Lunar Materials Technology Symposium

The meeting was organized around a possible lunar outpost scenario, featuring industrial technologies, systems, and components applicable to the extraction, processing, and fabrication of local materials. Acknowledged space resources experts as well as investigators from outside the field whose knowledge could be applied to space development activities were brought together. Presentations came from a variety of specialists in fields such as minerals processing, environmental control, and communications. The sessions of the symposium were divided into the following areas: resource characterization, energy management, materials processing, environment control, and automation and communications.

Source record↗
NASA NTRS1992

Shuttle entry guidance revisited

The Shuttle entry guidance concept is reviewed which is aimed at tracking a reference drag trajectory that leads to the specified range and velocity for the initiation of the terminal energy management phase. An approximate method of constructing the domain of attraction is proposed, and its validity is ascertained by simulation. An alternative guidance law yielding global exponential tracking in the absence of control saturation is derived using a feedback linearization method. It is noted that the alternative guidance law does not improve on the stability and performance of the current guidance law, for the operating domain and control capability of the Shuttle. It is suggested that the new guidance law with a larger operating domain and increased lift-to-drag capability would be superior.

Mease, Kenneth D.↗
NASA NTRS1993

Monte Carlo analysis of the Titan III/Transfer Orbit Stage guidance system for the Mars Observer mission

An important part of space launch vehicle mission planning for a planetary mission is the integrated analysis of guidance and performance dispersions for both booster and upper stage vehicles. For the Mars Observer mission, an integrated trajectory analysis was used to maximize the scientific payload and to minimize injection errors by optimizing the energy management of both vehicles. This was accomplished by designing the Titan III booster vehicle to inject into a hyperbolic departure plane, and the Transfer Orbit Stage (TOS) to correct any booster dispersions. An integrated Monte Carlo analysis of the performance and guidance dispersions of both vehicles provided sensitivities, an evaluation of their guidance schemes and an injection error covariance matrix. The polynomial guidance schemes used for the Titan III variable flight azimuth computations and the TOS solid rocket motor ignition time and burn direction derivations accounted for a wide variation of launch times, performance dispersions, and target conditions. The Mars Observer spacecraft was launched on 25 September 1992 on the Titan III/TOS vehicle. The post flight analysis indicated that a near perfect park orbit injection was achieved, followed by a trans-Mars injection with less than 2sigma errors.

Bell, Stephen C.↗
NASA NTRS1993

Numerical simulation of vehicle crashworthiness and occupant protection

Numerical simulation of vehicle crashworthiness and occupant protection are addressed. The vehicle crashworthiness design objectives are to design the vehicle structure for optimum impact energy absorption, and to design the restraint system (seatbelts, airbags, bolsters, etc.) for optimum occupant protection. The following approaches are taken; a major part of the impact energy is to be absorbed by the vehicle structure; the restraint components will provide protection against the remaining crash energy; certain vehicle components are designed to deform under specific types and speeds of impact in a desired mode for sound energy management; structural components such as front side rails, rear rails, door structure and pillars undergo large amounts of deformation; and with properly designed geometry and material these components assist in mitigating the effects of impact.

Saha, Nripen K.↗
NASA NTRS1994

The DQ Herculis stars

We review the properties of the DQ Herculis stars: cataclysmic variables containing an accreting, magnetic, rapidly rotating white dwarf. These stars are characterized by strong X-ray emission, high-excitation spectra, and very stable optical and X-ray pulsations in their light curves. There is considerable resemblance to their more famous cousins, the AM Herculis stars, but the latter class is additionally characterized by spin-orbit synchronism and the presence of strong circular polarization. We list eighteen stars passing muster as certain or very likely DQ Her stars. The rotational periods range from 33 s to 2.0 hr. Additional periods can result when the rotating searchlight illuminates other structures in the binary. A single hypothesis explains most of the observed properties: magnetically channeled accretion within a truncated disk. Some accretion flow still seems to proceed directly to the magnetosphere, however. The white dwarfs' magnetic moments are in the range 10(sup 32) - 10(sup 34) G cc, slightly weaker than in AM Her stars but with some probable overlap. The more important reason why DQ Hers have broken synchronism is probably their greater accretion rate and orbital separation. The observed L(sub x)/L(sub V) values are surprisingly low for a radially accreting white dwarf, suggesting that most of the accretion energy is not radiated in a strong shock above the magnetic pole. The fluxes can be more satisfactorily explained if most of the radial infall energy manages to bypass the shock and deposit itse lf directly in the white dwarf photosphere, where it should emerge as extreme ultraviolet (EUV) radiation. This also provides an adequate source of ionizing photons to power the high-excitation optical and UV emission lines. This is probably the DQ Her analog to the famous 'soft X-ray excess' in AM Her stars. However, unlike the AM Her case, this radiation has not been directly observed, so the analogy must not (yet) be embraced too firmly. There is some conventional wisdom today which segregates the short-period from the long-period DQ Her stars. But the observational grounds for this distinction are slim, except in one respect: X-ray emission from short-period systems appears to be weaker and softer. This must be due to the shallower depth of the potential well, and/or the greater difficulty the fast rotators have in enforcing radial accretion flow.

Patterson, Joseph↗
NASA NTRS1995

Integrated helmet mounted display concepts for air combat

A piloted simulation study was conducted in a dome simulator to evaluate several Helmet Mounted Display (HMD) formats developed as part of the NASA High Alpha Technology Program (HATP). The display formats conveyed energy management, spatial orientation, and weapons management information. The HMD format was compared to a generic Heads Up Display (HUD) typical of current operational fighter aircraft. Pilots were tasked to spend as much time in a weapon solution as possible, to have the correct weapon selected for the envelope they were in, and to avoid the adversary's weapon envelope as much as possible. Several different displays were tested individually and simultaneously to see how separate display concepts coexisted. Objective results showed that the ability for the pilot to select the correct weapon for the envelope he was in increased by 50% in a moderate workload condition and 90% in a high workload condition with the HMD format. In the post-test comments pilots generally favored the helmet display formats over the HUD formats with a few instances where pilots preferred a simple numeric readout of the parameter. Short term exposure effects of the HMD on visual acuity were also measured and showed no advers results.

Clark, Joseph W.↗
NASA NTRS1997

Retrofitting Laboratory Fume Hoods With Face Velocity Monitors at NASA Lewis Research Center

Extensive use and reliance on laboratory fume hoods exist at LeRC for the control of chemical hazards (nearly 175 fume hoods). Flow-measuring devices are necessary to continually monitor hood performance. The flow-measuring device should he tied into an energy management control system to detect problems at a central location without relying on the users to convey information of a problem. Compatibility concerns and limitations should always be considered when choosing the most effective flow-measuring device for a particular situation. Good practice on initial hood design and placement will provide a system for which a flow-measuring device may be used to its full potential and effectiveness.

Wagner, Ingrid E.↗
NASA NTRS1998

X-33 Attitude Control System Design for Ascent, Transition, and Entry Flight Regimes

The Vehicle Control Systems Team at Marshall Space Flight Center, Systems Dynamics Laboratory, Guidance and Control Systems Division is designing under a cooperative agreement with Lockheed Martin Skunkworks, the Ascent, Transition, and Entry flight attitude control system for the X-33 experimental vehicle. Ascent flight control begins at liftoff and ends at linear aerospike main engine cutoff (NECO) while Transition and Entry flight control begins at MECO and concludes at the terminal area energy management (TAEM) interface. TAEM occurs at approximately Mach 3.0. This task includes not only the design of the vehicle attitude control systems but also the development of requirements for attitude control system components and subsystems. The X-33 attitude control system design is challenged by a short design cycle, the design environment (Mach 0 to about Mach 15), and the X-33 incremental test philosophy. The X-33 design-to-launch cycle of less than 3 years requires a concurrent design approach while the test philosophy requires design adaptation to vehicle variations that are a function of Mach number and mission profile. The flight attitude control system must deal with the mixing of aerosurfaces, reaction control thrusters, and linear aerospike engine control effectors and handle parasitic effects such as vehicle flexibility and propellant sloshing from the uniquely shaped propellant tanks. The attitude control system design is, as usual, closely linked to many other subsystems and must deal with constraints and requirements from these subsystems.

Hall, Charles E.↗
NASA NTRS1998

Evaluation of GPS Coverage for the X-33 Michael-6 Trajectory

The onboard navigational system for the X-33 test flights will be based on the use of measurements collected from the Embedded Global Positioning System (GPS)/INS system. Some of the factors which will affect the quality of the GPS contribution to the navigational solution will be the number of pseudorange measurements collected at any instant in time, the distribution of the GPS satellites within the field of view, and the inherent noise level of the GPS receiver. The distribution of GPS satellites within the field of view of the receiver's antenna will depend on the receiver's position, the time of day, pointing direction of the antenna, and the effective cone angle of the antenna. The number of pseudorange measurements collected will depend upon these factors as well as the time required to lock onto a GPS satellite signal once the GPS satellite comes into the field of view of the antenna and the number of available receiver channels. The objective of this study is to evaluate the GPS coverage resulting from the proposed antenna pointing directions, the proposed antenna cone angles, and the effects due to the time of day for the X-33 Michael-6 trajectory from launch at Edwards AFB, California, to the start of the Terminal Area Energy Management (TAEM) phase on approach to Michael AAF, Utah.

Lundberg, John B.↗
NASA NTRS1999

Updated Results of Deterministic Reconfigurable Control Design for the X-33 Vehicle

In the event of a control surface failure, the purpose of a reconfigurable control system is to redistribute the control effort among the remaining working surfaces such that satisfactory stability and performance are retained. An Off-line Nonlinear General Constrained Optimization approach was used for the reconfigurable X-33 control design method. Three examples of failure are shown using a high fidelity 6 DOF simulation (case 1: ascent with a left body flap jammed at 25 deg.; case 2: entry with a right inboard elevon jam at 25 deg. and case 3: landing (TAEM) (Terminal Area Energy Management) with a left rudder jam at -30 deg.) Failure comparisons between responses with the nominal controller and reconfigurable controllers show the benefits of reconfiguration. Single jam aerosurface failures were considered, and failure detection and identification is considered accomplished in the actuator controller. The X-33 flight control system will incorporate reconfigurable flight control in the baseline system.

Cotting, M. Christopher↗
NASA NTRS1999

Entry Guidance for the Reusable Launch Vehicle

The X-33 Advanced Technology Demonstrator is a half-scale prototype developed to test the key technologies needed for a full-scale single-stage reusable launch vehicle (RLV). The X-33 is a suborbital vehicle that will be launched vertically, and land horizontally. The goals of this research were to develop an alternate entry guidance scheme for the X-33 in parallel to the actual X-33 entry guidance algorithms, provide comparative and complementary study, and identify potential new ways to improve entry guidance performance. Toward these goals, the nominal entry trajectory is defined by a piecewise linear drag-acceleration-versus-energy profile, which is in turn obtained by the solution of a semi-analytical parameter optimization problem. The closed-loop guidance is accomplished by tracking the nominal drag profile with primarily bank-angle modulation on-board. The bank-angle is commanded by a single full-envelope nonlinear trajectory control law. Near the end of the entry flight, the guidance logic is switched to heading control in order to meet strict conditions at the terminal area energy management interface. Two methods, one on ground-track control and the other on heading control, were proposed and examined for this phase of entry guidance where lateral control is emphasized. Trajectory dispersion studies were performed to evaluate the effectiveness of the entry guidance algorithms against a number of uncertainties including those in propulsion system, atmospheric properties, winds, aerodynamics, and propellant loading. Finally, a new trajectory-regulation method is introduced at the end as a promising precision entry guidance method. The guidance principle is very different and preliminary application in X-33 entry guidance simulation showed high precision that is difficult to achieve by existing methods.

Lu, Ping↗