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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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NASA NTRSDate not supplied

TPSAS-NF1676L-11021-DND

The Orion Pad Abort 1 Flight Test occurred on May 6th, 2010 at White Sands Missile Range in New Mexico. The test was the first for the Orion vehicle and was primarily a test of the Orion Launch Abort System’s ability to execute an abort from the launch pad. The test was successful and in most cases results were well within predicted outcomes. For this test, the thermal control subsystem was an inherent part of the flight operations team. The thermal team provided pre-launch predictions of component temperatures, set component temperature limits, and monitored those limits during the flight operations process as part of the primary flight operations team. The thermal team was also responsible for day-of-flight predictions of the main abort motor’s propellant mean bulk temperature, a key launch constraint for the test. After the test, thermal data was collected, reviewed, and compared against pre-launch predictions showing very good agreement in most areas. Based on the pre-test and flight operations experience, a detailed list of lessons learned has been developed and will be reviewed prior to potential future Orion or other such flight tests.

Joseph F Gasbarre↗
NASA NTRS1983

Application of in-flight thrust determination uncertainty

A numerical example is given of a previously proposed methodology for the evaluation of in-flight thrust measurement uncertainty, using data extracted from a performance report comparing two different missile prototypes under a variety of flight conditions. Attention is given to the data for the AGM-68B Air Launched Cruise Missile, which is powered by the F107 dual-spool, mixed flow turbofan engine. Assessments are made of the definition of the measurement process, instrumentation error estimation, the propagation of errors to thrust calculation, mathematical model errors, the in-flight thrust error component, and correction to standard conditions. It is concluded that in-flight thrust measurement uncertainty limits can be evaluated from measurement system error analysis results and test data for the missile evaluation process presently described.

Adams, G. R.↗
NASA NTRS2008

2007 Research and Engineering Annual Report

Selected research and technology activities at NASA Dryden Flight Research Center are summarized. These following activities exemplify the Center's varied and productive research efforts: Developing a Requirements Development Guide for an Automatic Ground Collision Avoidance System; Digital Terrain Data Compression and Rendering for Automatic Ground Collision Avoidance Systems; Nonlinear Flutter/Limit Cycle Oscillations Prediction Tool; Nonlinear System Identification Using Orthonormal Bases: Application to Aeroelastic/Aeroservoelastic Systems; Critical Aerodynamic Flow Feature Indicators: Towards Application with the Aerostructures Test Wing; Multidisciplinary Design, Analysis, and Optimization Tool Development Using a Genetic Algorithm; Structural Model Tuning Capability in an Object-Oriented Multidisciplinary Design, Analysis, and Optimization Tool; Extension of Ko Straight-Beam Displacement Theory to the Deformed Shape Predictions of Curved Structures; F-15B with Phoenix Missile and Pylon Assembly--Drag Force Estimation; Mass Property Testing of Phoenix Missile Hypersonic Testbed Hardware; ARMD Hypersonics Project Materials and Structures: Testing of Scramjet Thermal Protection System Concepts; High-Temperature Modal Survey of the Ruddervator Subcomponent Test Article; ARMD Hypersonics Project Materials and Structures: C/SiC Ruddervator Subcomponent Test and Analysis Task; Ground Vibration Testing and Model Correlation of the Phoenix Missile Hypersonic Testbed; Phoenix Missile Hypersonic Testbed: Performance Design and Analysis; Crew Exploration Vehicle Launch Abort System-Pad Abort-1 (PA-1) Flight Test; Testing the Orion (Crew Exploration Vehicle) Launch Abort System-Ascent Abort-1 (AA-1) Flight Test; SOFIA Flight-Test Flutter Prediction Methodology; SOFIA Closed-Door Aerodynamic Analyses; SOFIA Handling Qualities Evaluation for Closed-Door Operations; C-17 Support of IRAC Engine Model Development; Current Capabilities and Future Upgrade Plans of the C-17 Data Rack; Intelligent Data Mining Capabilities as Applied to Integrated Vehicle Health Management; STARS Flight Demonstration No. 2 IP Data Formatter; Space-Based Telemetry and Range Safety (STARS) Flight Demonstration No. 2 Range User Flight Test Results; Aerodynamic Effects of the Quiet Spike(tm) on an F-15B Aircraft; F-15 Intelligent Flight Controls-Increased Destabilization Failure; F-15 Integrated Resilient Aircraft Control (IRAC) Improved Adaptive Controller; Aeroelastic Analysis of the Ikhana/Fire Pod System; Ikhana: Western States Fire Missions Utilizing the Ames Research Center Fire Sensor; Ikhana: Fiber-Optic Wing Shape Sensors; Ikhana: ARTS III; SOFIA Closed-Door Flutter Envelope Flight Testing; F-15B Quiet Spike(TM) Aeroservoelastic Flight Test Data Analysis; and UAVSAR Platform Precision Autopilot Flight Results.

Stoliker, Patrick↗
NASA NTRS2016

Supersonic Flight Dynamics Test 2: Trajectory, Atmosphere, and Aerodynamics Reconstruction

The Supersonic Flight Dynamics Test is a full-scale flight test of aerodynamic decelerator technologies developed by the Low Density Supersonic Decelerator technology demonstration project. The purpose of the project is to develop and mature aerodynamic decelerator technologies for landing large-mass payloads on the surface of Mars. The technologies include a Supersonic Inflatable Aerodynamic Decelerator and supersonic parachutes. The first Supersonic Flight Dynamics Test occurred on June 28th, 2014 at the Pacific Missile Range Facility. The purpose of this test was to validate the test architecture for future tests. The flight was a success and, in addition, was able to acquire data on the aerodynamic performance of the supersonic inflatable decelerator. The Supersonic Disksail parachute developed a tear during deployment. The second flight test occurred on June 8th, 2015, and incorporated a Supersonic Ringsail parachute which was redesigned based on data from the first flight. Again, the inflatable decelerator functioned as predicted but the parachute was damaged during deployment. This paper describes the instrumentation, analysis techniques, and acquired flight test data utilized to reconstruct the vehicle trajectory, main motor thrust, atmosphere, and aerodynamics.

Karlgaard, Christopher D.↗
NASA NTRS1991

An overview of autonomous rendezvous and docking system technology development at General Dynamics

The Centaur avionics suite is undergoing a dramatic modernization for the commercial, DoD Atlas and Titan programs. The system has been upgraded to the current state-of-the-art in ring laser gyro inertial sensors and Mil-Std-1750A processor technology. The Cruise Missile avionic system has similarly been evolving for many years. Integration of GPS into both systems has been underway for over five years with a follow-on cruise missile system currently in flight test. Rendezvous and Docking related studies have been conducted for over five years in support of OMV, CTV, and Advanced Upper Stages, as well as several other internal IR&D's. The avionics system and AR&D simulator demonstrated to the SATWG in November 1990 has been upgraded considerably under two IR&D programs in 1991. The Centaur modern avionics system is being flown in block upgrades which started in July of 1990. The Inertial Navigation Unit will fly in November of 1991. The Cruise Missile avionics systems have been fully tested and operationally validated in combat. The integrated AR&D system for space vehicle applications has been under development and testing since 1990. A Joint NASA / GD ARD&L System Test Program is currently being planned to validate several aspects of system performance in three different NASA test facilities in 1992.

Kuenzel, Fred↗
NASA NTRS1979

Bumblebee program, aerodynamic data. Part 2: Flow fields at Mach number 2.0

Available flow field data which can be used in validating theoretical procedures for computing flow fields around supersonic missiles are presented. Tabulated test data are given which define the flow field around a conical-nosed cylindrical body in a crossflow plane corresponding to a likely tail location. The data were obtained at a Mach number of 2.0 for an angle of attack of 0 to 23 degrees. The data define the flow field for cases both with and without a forward wing present.

Barnes, G. A.↗
NASA NTRS1992

Aeroelastic stability analyses of two counter rotating propfan designs for a cruise missile model

Aeroelastic stability analyses were performed to insure structural integrity of two counterrotating propfan blade designs for a NAVY/Air Force/NASA cruise missile model wind tunnel test. This analysis predicted if the propfan designs would be flutter free at the operating conditions of the wind tunnel test. Calculated stability results are presented for the two blade designs with rotational speed and Mach number as the parameters. A aeroelastic analysis code ASTROP2 (Aeroelastic Stability and Response of Propulsion Systems - 2 Dimensional Analysis), developed at LeRC, was used in this project. The aeroelastic analysis is a modal method and uses the combination of a finite element structural model and two dimensional steady and unsteady cascade aerodynamic models. This code was developed to analyze single rotation propfans but was modified and applied to counterrotating propfans for the present work. Modifications were made to transform the geometry and rotation of the aft rotor to the same reference frame as the forward rotor, to input a non-uniform inflow into the rotor being analyzed, and to automatically converge to the least stable aeroelastic mode.

Mahajan, Aparajit J.↗
NASA NTRS1995

GPS as an orbit determination subsystems

This paper evaluates the use of Global Positioning System (GPS) receivers as a primary source of tracking data for low-Earth orbit satellites. GPS data is an alternative to using range, azimuth, elevation, and range-rate (RAER) data from the Air Force Satellite Control Network antennas, the Space Ground Link System (SGLS). This evaluation is applicable to missions such as Skipper, a joint U.S. and Russian atmosphere research mission, that will rely on a GPS receiver as a primary tracking data source. The Detachment 2, Space and Missile Systems Center's Test Support Complex (TSC) conducted the evaluation based on receiver data from the Space Test Experiment Platform Mission O (STEP-O) and Advanced Photovoltaic and Electronics Experiments (APEX) satellites. The TSC performed orbit reconstruction and prediction on the STEP-0 and APEX vehicles using GPS receiver navigation solution data, SGLS RAER data, and SGLS anglesonly (azimuth and elevation) data. For the STEP-O case, the navigation solution based orbits proved to be more accurate than SGLS RAER based orbits. For the APEX case, navigation solution based orbits proved to be less accurate than SGLS RAER based orbits for orbit prediction, and results for orbit reconstruction were inconclusive due to the lack of a precise truth orbit. After evaluating several different GPS data processing methods, the TSC concluded that using GPS navigation solution data is a viable alternative to using SGLS RAER data.

Fennessey, Richard↗
NASA NTRS1997

NiH2 Battery Reconditioning for LEO Applications

This paper summarizes reasons for and benefits of reconditioning nickel-hydrogen (NiH2) batteries used for Low Earth Orbit (LEO) applications. NiH2 battery cells do not have the classic discharge voltage problems more commonly associated with nickel-cadmium (NiCd) cells. This is due, in part, to use of hydrogen electrodes in place of cadmium electrodes. The nickel electrode, however, does have a similar discharge voltage signature for both cell designs. This can have an impact on LEO applications where peak loads at higher relative depths of discharge can impact operations. Periodic reconditioning provides information which can be used for analyzing long term performance trends to predict usable capacity to a specified voltage level. The reconditioning process described herein involves discharging NiH2 batteries at C/20 rates or less, to an average cell voltage of 1.0 volts or less. Recharge is performed at nominal C/5 rates to specified voltage/temperature (V/T) charge levels selected to restore required capacity with minimal overcharge. Reconditioning is a process of restoring reserve capacity lost on cycling, which is commonly called the memory effect in NiCd cells. This effect is characterized by decreases in the discharge voltage curve with operational life and cycling. The end effect of reconditioning NiH2 cells may be hidden in the versatility, of that design over the NiCd cell design and its associated negative electrode fading problem. The process of deep discharge at lower rates by way of reconditioning tends to redistribute electrolyte and water in the NiH2 cell electrode stack, while improving utilization and charge efficiency. NiH2 battery reconditioning effects on life are considered beneficial and may, in fact. extend life based on NiCd experience. In any case, usable capacity data obtained from reconditioning is required for performance evaluation and trend analysis. Characterization and life tests have provided the historical data base used to determine the need for reconditioning in most battery applications. The following sections briefly describe the background of NiH2 battery reconditioning and testing at Lockheed Martin Missiles & Space (LMMS) and other aerospace companies.

Armantrout, J. D.↗
NASA NTRS1955

Development of a high-velocity free-flight launcher : the Ames light-gas gun

Recent interest in long-range missiles has stimulated a search for new experimental techniques which can reproduce in the laboratory the high temperatures and Mach numbers associated with the missiles' flight. One promising possibility lies in free-flight testing of laboratory models which are flown at the full velocity of the missile. In this type of test, temperatures are approximated and aerodynamic heating of the model is representative of that experienced by the missile in high-velocity flight. A prime requirement of the free-flight test technique is a device which had the capacity for launching models at the velocities desired. In response to thie need, a gun firing light models at velocities up to 15,000 feet per second has been developed at the Ames Aeronautical Laboratory. The design of this gun, the analysis of its performance, and the results of the initial firing trials are described in this paper. The firing trials showed that the measured velocities and pressures agreed well with the predicted values. Also, the erosion of the launch tube was very small for the eleven rounds fired. The performance of the gun suggests that it will prove to be a satisfactory launcher for high-velocity free-flight tests. However, it should be mentioned that only the gross performance has been evaluated so far, and, consequently, the operation of the gun must be investigated in further detail before its performance can be reliably predicted over its full operating range.

Charters, A C↗
NASA NTRS1959

An Investigation of a 0.05-Scale Model of the XSM-64A Navaho Missile and Booster: Force Study at Mach Numbers from 1.77 to 3.51 - Part 1

An investigation has been conducted in the Langley Unitary Plan wind tunnel to determine the aerodynamic loads and the static longitudinal and lateral stability of a 0.05-scale model of the XSM-64A Navaho missile and booster and its various components. Tests were conducted through a Mach number range of 1.77 to 3.51 with a corresponding Reynolds number range of 2.4 x 10(exp 6) to 2.9 x 10(exp 6). Results are presented for an angle-of-attack range of -8 deg to 4 deg for the missile-booster combination and -10 deg to 10 deg for the missile-alone configuration. Tests for both configurations were conducted through an angle-of-sideslip range of -8 deg to 8 deg. Also presented are some effects on the model characteristics of the deflection of various components including canard, tip aileron, vertical stabilizer, speed brakes, and booster pitch and yaw thrust chambers. The various components on which loads were measured include the wing, tip aileron, rudder, booster, booster separating surface, booster fin, and booster yaw and pitch thrust chambers. These data are presented without analysis.

Church, James D.↗
NASA NTRS1986

Effects of some geometric variations on missile aerodynamic characteristics at supersonic speeds

A summary of some results from tests of a wing-body general research missile model is presented for a Mach number range up to 4.63. A basic ogive-cylinder body with a length-to-diameter ratio of 10 was used to which was attached a series of wing planforms. The planforms included a family of delta wings and a family of rectangular wings having a constant root chord but varying spans so that wings of constant exposed area could be compared. In addition, a cranked-tip planform was included and a rectangular planform with reduced chord. Some results are presented for wing-body-tail configurations - one utilizing a cranked wing planform and one with wings having a constant root chord and span, but tip chords that were 0, 20, and 40 percent of the root chord.

Spearman, M. L.↗
NASA NTRS1986

Effects of some geometric variations on missile aerodynamic characteristics at supersonic speeds

A summary of some results from tests of a wing-body general research missile model is presented for a Mach number range up to 4.63. A basic ogive-cylinder body with a length-to-diameter ratio of 10 was used to which was attached a series of wing planforms. The planforms included a family of delta wings and a family of rectangular wings having a constant root chord but varying spans so that wings of constant exposed area could be compared. In addition, a cranked-tip planform was included and a rectangular planform with reduced chord. Some results are presented for wing-body-tail configurations - one utilizing a cranked wing planform and one with wings having a constant root chord and span, but tip chords that were 0, 20, and 40 percent of the root chord.

Spearman, M. L.↗
NASA NTRS1999

Pyroshock Environments Characterized for Spacecraft Missions

Pyrotechnic shock, or pyroshock, is the transient response of a structure to loading induced by the ignition of pyrotechnic (explosive or propellant activated) devices. These devices are typically used to separate structural systems (e.g., separate a spacecraft from a launch vehicle) and deploy appendages (e.g., solar panels). Pyroshocks are characterized by high peak acceleration, high-frequency content, and short duration. Because of their high acceleration and high-frequency, pyroshocks can cause spaceflight hardware to fail. Verifying by test that spaceflight hardware can withstand the anticipated shock environment is considered essential to mission success. The Earth Observing System (EOS) AM-1 spacecraft for NASA's Mission to Planet Earth is scheduled to be launched on an Atlas IIAS vehicle in 1999, and the NASA Lewis Research Center is the launch vehicle integrator for this NASA Goddard Space Flight Center spacecraft. The EOS spacecraft was subjected to numerous ground shock tests to verify that its scientific instruments and avionics components will withstand the shock-induced vibration produced when the spacecraft separates from the launch vehicle. Shock test data from these tests represent the third largest available pyroshock database in the United States. Future spacecraft missions will directly benefit from the knowledge gained from these tests. The payload separation system used for EOS is a new system that operates by firing six separation nuts. This system was tested to verify its functional operation and to characterize the resulting shock levels. The launch vehicle contractor (Lockheed Martin Astronautics) and spacecraft contractor (Lockheed Martin Missiles & Space) completed 16 separation test firings. This resulted in an unusually large amount of pyroshock data. Typically, only one or two pyroshock test firings are performed for a spacecraft mission. Because of the size of this separation system shock database, engineers were able to perform unique statistical analyses to characterize the distribution of the test data. For example, it was proven that the shock data follow a lognormal distribution, a concept often assumed but rarely proven. The test-to-test repeatability of the shock source level was analyzed, and the effects of various test configurations and separation nut production lots were examined and quantified. Engineers investigated the change in shock level as the shock traveled from the spacecraft separation interface to the avionics components of the upper stage and analyzed the effects of the structural fidelity (simulator versus real) of the components and their weight on vibrational response. In addition, the shock attenuation with distance and across joints was quantified and compared with concepts originally generated in 1970, and the effects of separation nut preload and firing sequences effects were examined. Because of this EOS shock testing and the analyses performed at NASA Lewis, a significant amount of new information on pyroshock and its characteristics is now available to the aerospace industry. We hope that this information will help future spacecraft test planners to perform better and cheaper spacecraft separation shock tests and to better understand their test data.

Hughes, William O.↗
NASA NTRS1955

Longitudinal Stability and Control Characteristics as Determined by the Rocket-model Technique for an Inline, Cruciform, Canard Missile Configuration with a Low-aspect-ratio Wing Having Trailing-edge Flap Controls for a Mach Number Range of 0.7 to 1.

Two full-scale models of an inline, cruciform, canard missile configuration having a low-aspect-ratio wing equipped with flap-type controls were flight tested in order to determine the missile's longitudinal aerodynamic characteristics. Stability derivatives and control and drag characteristics are presented for a range of Mach number from 0.7 to 1.8. Nonlinear lift and moment curves were noted for the angle - of-attack range of this test (0 deg to 8 deg). The aerodynamic-center location for angles of attack near 50 remained nearly constant for supersonic speeds at 13.5 percent of the mean aerodynamic chord; whereas for angles of attack near 0 deg, there was a rapid forward movement of the aerodynamic center as the Mach number increased. At a control deflection of 0 deg, the missile's response to the longitudinal control was in an essentially fixed space plane which was not coincident with the pitch plane as a result of the missile rolling. As a consequence, stability characteristics were determined from the resultant of pitch and yaw motions. The damping-in-pitch derivatives for the two angle -of-attack ranges of the test are in close agreement and varied only slightly with Mach number. The horn-balanced trailing-edge flap was effective in producing angle of attack over the Mach number range.

Baber, Hal T , Jr↗
NASA NTRS1955

Longitudinal Stability and Control Characteristics as Determined by the Rocket-Model Technique for an Inline, Cruciform, Canard Missile Configuration with a Low-Aspect-Ratio Wing Having Trailing-Edge Flap Controls for a Mach Number Range of 0.7 to 1.8

Two full-scale models of an inline, cruciform, canard missile configuration having a low-aspect-ratio wing equipped with flap-type controls were flight tested in order to determine the missile's longitudinal aerodynamic characteristics. Stability derivatives and control and drag characteristics are presented for a range of Mach number from 0.7 to 1.8. Nonlinear lift and moment curves were noted for the angle-of-attack range of this test (0 deg to 8 deg ). The aerodynamic-center location for angles of attack near 5 deg remained nearly constant for supersonic speeds at 13.5 percent of the mean aerodynamic chord; whereas for angles of attack near O deg, there was a rapid forward movement of the aerodynamic center as the Mach number increased. At a control deflection of O deg, the missile's response to the longitudinal control was in an essentially fixed space plane which was not coincident with the pitch plane as a result of the missile rolling. As a consequence, stability characteristics were determined from the resultant of pitch and yaw motions. The damping-in-pitch derivatives for the two angle-of-attack ranges of the test are in close agreement and varied only slightly with Mach number. The horn-balanced trailing-edge flap was effective in producing angle of attack over the Mach number range.

Baber, H. T., Jr.↗