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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 613 records · Page 34

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NASA NTRS1970

Polynomial filter estimation of range and range rate for terminal rendezvous

A study was made of a polynomial filter for computing range rate information from CSM VHF range data. The filter's performance during the terminal phase of the rendezvous is discussed. Two modifications of the filter were also made and tested. A manual terminal rendezvous was simulated and desired accuracies were achieved for vehicles on an intercept trajectory, except for short periods following each braking maneuver when the estimated range rate was initially in error by the magnitude of the burn.

Philips, R.↗

NASA NTRS1972

Application of an approach and landing system model to the space shuttle orbiter vehicle

An analytic model and procedure are described which can be used to estimate probability distributions of touchdown conditions, without utilizing a Monte Carlo simulation. Example applications are presented, including one-step and two-step flare strategies for the space shuttle orbiter. The computed probability distributions for the two-step flare are compared with the corresponding histograms from an independent Monte Carlo simulation of the same situation. The computed distributions are very good fits to the Monte-Carlo-generated histograms for pertinent longitudinal and lateral variables at three points (including touchdown) along the vehicle's final approach trajectory. The model, thus verified, has more inherent power for assessing the effects of system element changes than conventional Monte Carlo techniques.

Johnson, W. A.↗

NASA NTRS1973

S-band transponder experiment

The purpose of this experiment was to measure the variations in the lunar gravitational field near the trajectory of orbiting space vehicles (the command and service module (CSM) and the small particles and fields subsatellites ejected from the Apollo 15 and 16 spacecraft). New information has been obtained from all Apollo orbiting spacecraft; however, this report shall be limited to the results from the Apollo 17 CSM and the Apollo 16 subsatellite. The data acquired are precise speed measurements of the orbiting spacecraft from which accelerations or gravity profiles may be inferred. Feature resolution is controlled by the spacecraft altitude and is almost a direct relationship (i.e., data taken from a 50-km altitude will resolve approximately a 50-km feature). Therefore, revolutions 3 to 12, when the CSM was in the low-altitude orbits, provided the clearest information.

Sjogren, W. L.↗

NASA NTRS1976

Continuation of advanced crew procedures development techniques

An operational computer program, the Procedures and Performance Program (PPP) which operates in conjunction with the Phase I Shuttle Procedures Simulator to provide a procedures recording and crew/vehicle performance monitoring capability was developed. A technical synopsis of each task resulting in the development of the Procedures and Performance Program is provided. Conclusions and recommendations for action leading to the improvements in production of crew procedures development and crew training support are included. The PPP provides real-time CRT displays and post-run hardcopy output of procedures, difference procedures, performance data, parametric analysis data, and training script/training status data. During post-run, the program is designed to support evaluation through the reconstruction of displays to any point in time. A permanent record of the simulation exercise can be obtained via hardcopy output of the display data and via transfer to the Generalized Documentation Processor (GDP). Reference procedures data may be transferred from the GDP to the PPP. Interface is provided with the all digital trajectory program, the Space Vehicle Dynamics Simulator (SVDS) to support initial procedures timeline development.

Arbet, J. D.↗

NASA NTRS1986

Libration-point staging concepts for Earth-Mars transportation

The use of libration points as transfer nodes for an Earth-Mars transportation system is briefly described. It is assumed that a reusable Interplanetary Shuttle Vehicle (ISV) operates between the libration point and Mars orbit. Propellant for the round-trip journey to Mars and other supplies would be carried from low Earth orbit (LEO) to the ISV by additional shuttle vehicles. Different types of trajectories between LEO and libration points are presented, and approximate delta-V estimates for these transfers are given. The possible use of lunar gravity-assist maneuvers is also discussed.

Robert Farquhar↗

NASA NTRS1990

Reacting flows

Features of chemically reacting separated hypersonic flows are identified and issues concerning their analysis and simulation are discussed. Emphasis is placed on flows of high temperature dissociating and ionizing air and current methods for studying and characterizing these flows, including separation, are reviewed. The aeroassist orbital transfer vehicle and its flight trajectory are used for illustration. Thermochemical nonequilibrium phenomena are emphasized and extension of continuum analysis to the high altitude slip-flow regime is considered.

Deiwert, George S.↗

NASA NTRS1990

Optimal plane change by low aerodynamic forces

This paper presents the exact dimensionless equations of motion and the necessary conditions for the computation of the optimal trajectories of a hypervelocity vehicle flying through a nonrotating spherical planetary atmosphere. It is shown that there are two types of maneuvers with nearly identical plane change. In the hard maneuver, the vehicle is pulled down to low altitude for aerodyamic plane change before exit at the prescribed final speed. In the slow maneuver which is described in detail in this paper, the vehicle remains in orbital flight with a small incremental plane change during each passage through the perigee. This maneuver requires several revolutions, and the technique for computation is similar to that in the problem of contraction of orbit.

Vinh, Nguyen X.↗

NASA NTRS1991

Lunar base siting

As with any planetary body, the lunar surface is quite heterogeneous. There are widely dispersed sites of particular interest for known and potential resource availability, selenology, and lunar observatories. Discriminating characteristics include solar illumination, view of earth, local topography, engineering properties of the regolith and certain geological features, and local mineralogy and petrology. Space vehicle arrival and departure trajectories constitute a minor consideration. Over time, a variety of base sites will be developed serving different purposes. Resource-driven sites may see the fastest growth during the first decades of lunar development, but selection of the most favorable sites is likely to be driven by suitability for a combination of activities. As on earth, later development may be driven by geographical advantages of surface transportation routes. With the availability of near-constant sunlight for power generation, as well as permanently shadowed areas at cryogenic temperatures, polar sites are attractive because they require substantially less earth-launched mass and lower equipment complexity for an initial permanent base. Discovery of accessible volatiles reservoirs, either in the form of polar permafrost or gas reservoirs at other locations, would dramatically increase the attractiveness of any site from a logistical support and selenological point of view. Amid such speculation, no reliable evidence of such volatiles exist. More reliable evidence exists for areas of certain mineral concentrations, such as ilmenite, which could form a feedstock for some proposed resource extraction schemes. While tentative selections of advantageous base sites are made, new data from lunar polar orbiters and the Galileo polar flybys would be very helpful.

Staehle, Robert L.↗

NASA NTRS1990

Optimal multiple-pass aeroassisted plane change

This paper presents the exact dimensionless equation of motion and the necessary conditions for the computation of the optimal trajectories of a hypervelocity vehicle flying through a non-rotating spherical planetary atmosphere. Numerical solution is then presented for the case when the vehicle makes several passages through the atmosphere near the perigee of its orbit. While the orbit is slowly contracting, aerodynamic maneuver is performed to obtain the maximum plane change. Several plots were presented to show the optimal variations of the lift coefficient and the bank angle and the various elements of the orbit.

Vinh, Nguyen X.↗

NASA NTRS1991

Control algorithms for aerobraking in the Martian atmosphere

The Analytic Predictor Corrector (APC) and Energy Controller (EC) atmospheric guidance concepts were adapted to control an interplanetary vehicle aerobraking in the Martian atmosphere. Changes are made to the APC to improve its robustness to density variations. These changes include adaptation of a new exit phase algorithm, an adaptive transition velocity to initiate the exit phase, refinement of the reference dynamic pressure calculation and two improved density estimation techniques. The modified controller with the hybrid density estimation technique is called the Mars Hybrid Predictor Corrector (MHPC), while the modified controller with a polynomial density estimator is called the Mars Predictor Corrector (MPC). A Lyapunov Steepest Descent Controller (LSDC) is adapted to control the vehicle. The LSDC lacked robustness, so a Lyapunov tracking exit phase algorithm is developed to guide the vehicle along a reference trajectory. This algorithm, when using the hybrid density estimation technique to define the reference path, is called the Lyapunov Hybrid Tracking Controller (LHTC). With the polynomial density estimator used to define the reference trajectory, the algorithm is called the Lyapunov Tracking Controller (LTC). These four new controllers are tested using a six degree of freedom computer simulation to evaluate their robustness. The MHPC, MPC, LHTC, and LTC show dramatic improvements in robustness over the APC and EC.

Ward, Donald T.↗

NASA NTRS1993

Investigation of the launch pad abort capabilities of the HL-20 lifting body

The capability of the HL-20 lifting body spacecraft to perform an abort maneuver from the launch pad to a horizontal landing was studied at NASA Langley Research Center. This study involved both piloted and batch simulation models of the vehicle. A point-mass model of the vehicle was used for trajectory optimization studies. The piloted simulation was performed in the Visual Motion Simulator in fixed-base mode. A candidate maneuver was developed and refined for the worst-case launch-pad-to-landing-site geometry using an iterative procedure of off-line maneuver analysis followed by piloted evaluations and heuristic improvements to the candidate maneuver. The resulting maneuver demonstrates the launch site abort capability of the HL-20 and dictates requirements for nominal abort motor performance. The sensitivity of the maneuver to variations in several design parameters was documented.

Jackson, E. B.↗

NASA NTRS1992

Robust control algorithms for Mars aerobraking

Four atmospheric guidance concepts have been adapted to control an interplanetary vehicle aerobraking in the Martian atmosphere. The first two offer improvements to the Analytic Predictor Corrector (APC) to increase its robustness to density variations. The second two are variations of a new Liapunov tracking exit phase algorithm, developed to guide the vehicle along a reference trajectory. These four new controllers are tested using a six degree of freedom computer simulation to evaluate their robustness. MARSGRAM is used to develop realistic atmospheres for the study. When square wave density pulses perturb the atmosphere all four controllers are successful. The algorithms are tested against atmospheres where the inbound and outbound density functions are different. Square wave density pulses are again used, but only for the outbound leg of the trajectory. Additionally, sine waves are used to perturb the density function. The new algorithms are found to be more robust than any previously tested and a Liapunov controller is selected as the most robust control algorithm overall examined.

Shipley, Buford W., Jr.↗

NASA NTRS1994

Launch-pad abort capabilities of the HL-20 lifting body

The capability of the HL-20 lifting body to perform an abort maneuver from the launch pad to a horizontal landing was studied. The study involved both piloted and batch simulation models of the vehicle. A point-mass model of the vehicle was used for trajectory optimization studies. The piloted simulation was performed in the Langley Visual/Motion Simulator in the fixed-base mode. A candidate maneuver was developed and refined for the worst-case launch-pad-to-landing-site geometry with an iterative procedure of off-line maneuver analysis followed by piloted evaluations and heuristic improvements to the candidate maneuver. The resulting maneuver demonstrates the launch-site abort capability of the HL-20 and dictates requirements for nominal abort-motor performance. The sensitivity of the maneuver to variations in several design parameters was documented.

Jackson, E. Bruce↗

NASA NTRS1994

Matlab as a robust control design tool

This presentation introduces Matlab as a tool used in flight control research. The example used to illustrate some of the capabilities of this software is a robust controller designed for a single stage to orbit air breathing vehicles's ascent to orbit. The global requirements of the controller are to stabilize the vehicle and follow a trajectory in the presence of atmospheric disturbances and strong dynamic coupling between airframe and propulsion.

Gregory, Irene M.↗

NASA NTRS1996

Pulse Detonation Engine Air Induction System Analysis

A preliminary mixed-compression inlet design concept for potential pulse-detonation engine (PDE) powered supersonic aircraft was defined and analyzed. The objectives of this research were to conceptually design and integrate an inlet/PDE propulsion system into a supersonic aircraft, perform time-dependent CFD analysis of the inlet flowfield, and to estimate the installed PDE cycle performance. The study was baselined to a NASA Mach 5 Waverider study vehicle in which the baseline over/under turboramjet engines were replaced with a single flowpath PDE propulsion system. As much commonality as possible was maintained with the baseline configuration, including the engine location and forebody lines. Modifications were made to the inlet system's external ramp angles and a rotating cowl lip was incorporated to improve off-design inlet operability and performance. Engines were sized to match the baseline vehicle study's ascent trajectory thrust requirement at Mach 1.2. The majority of this study was focused on a flight Mach number of 3.0. The time-dependent Navier Stokes CFD analyses of a two-dimensional approximation of the inlet was conducted for the Mach 3.0 condition. The Lockheed Martin Tactical Aircraft Systems-developed FALCON CFD code with a two equation 'k-1' turbulence model was used. The downstream PDE was simulated by an array of four sonic nozzles in which the flow areas were rapidly varied in various opening/closing combinations. Results of the CFD study indicated that the inlet design concept operated successfully at the Mach 3.0 condition, satisfying mass capture, total pressure recovery, and operability requirements. Time-dependent analysis indicated that pressure and expansion waves from the simulated valve perturbations did not effect the inlet's operability or performance.

Pegg, R. J.↗

NASA NTRS1998

X-34 Experimental Aeroheating at Mach 6 and 10

Critical technologies are being developed to support the goals of the NASA Office of Aeronautics and Space Transportation Technology Access to Space initiative for next-generation reusable space transportation systems. From the perspective of aerothermodynamic performance throughout the flight trajectory, the Reusable Launch Vehicle program incorporates conceptual analysis, ground-based testing, and computational fluid dynamics to provide flyable suborbital flight demonstrator vehicles. This report provides an overview of the hypersonic aeroheating wind tunnel test program conducted at the NASA Langley Research Center in support of one of these vehicles, the X-34 small reusable technology demonstrator program. Global surface heat transfer images, surface streamline patterns, and shock shapes were measured on 0.0153- and 0.0183-scale models of proposed X-34 flight vehicles at Mach 6 and 10 in air. The primary parametrics that were investigated include angles-of-attack from 0 to 35 deg. and freestream unit Reynolds numbers from 0.5 to 8 million per foot (which was sufficient to produce laminar, transitional, and turbulent heating data), both with and without control surface deflections. Comparisons of the experimental data to computational predictions are included, along with a discussion of the implications of some of the experimental flow features for the flight vehicle.

Berry, Scott A.↗

NASA NTRS1995

Capsule Aerothermodynamics. CFD Prediction Methodology; Database for CFD Validation; Capsule Flight Programs: AFE and Apollo Lessons Learned

Deiwert, George S.↗

NASA NTRS1960

Experience with a three-axis side-located controller during a static and centrifuge simulation of the piloted launch of a manned multistage vehicle

An investigation was conducted to determine a human pilot's ability to control a multistage vehicle through the launch trajectory. The simulation was performed statically and dynamically by utilizing a human centrifuge. An interesting byproduct of the program was the three-axis side-located controller incorporated for pilot control inputs. This method of control proved to be acceptable for the successful completion of the tracking task during the simulation. There was no apparent effect of acceleration on the mechanical operation of the controller, but the pilot's control feel deteriorated as his dexterity decreased at high levels of acceleration. The application of control in a specific control mode was not difficult. However, coordination of more than one mode was difficult, and, in many instances, resulted in inadvertent control inputs. The acceptable control harmony at an acceleration level of 1 g became unacceptable at higher acceleration levels. Proper control-force harmony for a particular control task appears to be more critical for a three-axis controller than for conventional controllers. During simulations in which the pilot wore a pressure suit, the nature of the suit gloves further aggravated this condition.

Andrews, William H.↗