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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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The possible uses of pulsed carbon dioxide lasers for analysis of plasmas and flows need appropriate infrared image converters. Emphasis was placed on liquid crystal detectors and their operational modes. Performance characterstics and selection criteria, such as high sensitivity, short reaction time, and high spatial resolution are discussed.
The Performance Verification Test Procedure is given for the Halogen Occultation Experiment (HALOE) instrument, which is being developed in house at the Langley Research Center for the Upper Atmosphere Research Satellite (UARS). This procedure is used for comprehensive performance testing of the HALOE instrument which occurs before, during, and after flight environmental tests. The radiometric performance tests include noise, drift, linearity, instantaneous field-of-view, cal wheel gas cell characterization, and self thermal emissions. Pointer/tracker performance tests include sun sensor performance, gimbal performance, control system performance, and boresight alignment. In addition, the instrument is tested functionally in simulated orbit sequences and all command operating modes are exercised. The data analysis required for each test is specified and pass/fail criteria are given where applicable. This test will fully demonstrate the HALOE instrument's ability to achieve science mission requirements. The HALOE instrument is a gas correlation radiometer that measures vertical distribution of eight upper atmospheric constituents: O3, HCl, HF, NO, CH4, H2O, NO2, and CO2.
As part of the RL10 Rocket Engine Product Improvement Program, Oxidizer Heat Exchanger (OHE) stages 1, 2, and 3 were designed and fabricated during late 1983 and early 1984. The purpose of the OHE is to provide gaseous oxygen to the propellant injector for stable engine operation at tank head idle and pumped idle operating modes. This report summarizes the OHE stages 1 and 3 rig testing, and includes the separation of the stage 1-and-2 assembly and the remanifolding of stage 1. The OHE performance analysis and analytical model modifications for both stages are also presented. The flow tests were accomplished during the time period from 9 October 1984 to 12 November 1984.
The validation of the guidance, navigation, and control system of the Space Shuttle is explained. The functions of the ascent, on-board, and entry mission phases software of the navigation system are described. The common facility testing, which evaluates the simulations to be used in the navigation validation, is examined. The standard preflight analysis of the operational modes of the navigation software and the post-flight navigation analysis are explained. The conversion of the data into a useful reference frame and the use of orbit parameters in the analysis of the data are discussed. Upon entry the data received are converted to flags, ratios, and residuals in order to evaluate performance and detect errors. Various programs developed to support navigation validation are explained. A number of events that occurred with the Space Shuttle's navigation system are described.
This paper describes the Orbital Maneuvering Vehicle (OMV) concept and its intended role. It recaps the past activities leading up to the current concept and summarizes the present status and plans. The various types of missions, operating modes, and performance capability are described. Typical mission scenarios for servicing missions from both the Shuttle and the Space Station are described.
This paper describes the Orbital Maneuvering Vehicle (OMV) concept and its intended role. It recaps the past activities leading up to the current concept and summarizes the present status and plans. The various types of missions and operating modes required by the OMV are described as the basis of the guidance, navigation and control (GN&C) requirements. The general GN&C problem is outlined with potential hardware solutions.
High-pressure annular seal leakage and dynamic coefficients vary with eccentricity and misalignment. Recent seal leakage data with both concentric and fully eccentric alignments support the seal leakage model with surface roughness and eccentricity effects included. In this paper, the seal dynamic coefficient calculation has been generalized and allows direct calculation of the seal dynamic coefficients at any circumferential location. The generalized solution agrees with the results obtained by using the calculated values of an earlier paper and performing a coordinate transformation. The analysis results coincide with the measured data in showing that the stiffness and damping matrices of seal coefficients are not skew symmetric, and the main diagonal seal coefficients are not equal. The measured direct stiffnesses were found higher than predicted by the concentric seal theory, but this may be explained by the presence of eccentricity in the test operating mode.
Preliminary performance goals for the Space Telescope Imaging Spectrograph (STIS), selected by NASA for definition study as a second generation instrument on the Hubble Space Telescope (HST) are outlined. The STIS is a 2-dimensional spectrograph that will operate from 1050A-11000A at the limiting HST spatial resolution of 0.05 arcsec FWHM, with spectral resolutions of 100, 1200, 20,000, and 100,000 and a maximum field-of-view of 50X50 arcsec. Its basic operating modes include echelle mode (like IUE), long slit mode (for simultaneous spectra at 1000 positions on the target), slitless spectrograph mode, coronographic spectroscopy, photon time-tagging, and direct imaging.
The RL10-2B engine, a derivative of the RL10, is capable of multimode thrust operation. This engine operates at two low thrust levels: tank head idle (THI), which is approximately 1 to 2% of full thrust, and pumped idle (PI), which is 10% of full thrust. Operation at THI provides vehicle propellant settling thrust and efficient engine thermal conditioning; PI operation provides vehicle tank pre-pressurization and maneuver thrust for low-g deployment. Stable combustion of the RL10-2B engine during the low thrust operating modes can be accomplished by using a heat exchanger to supply gaseous oxygen to the propellant injector. The oxidizer heat exchanger (OHE) vaporizes the liquid oxygen using hydrogen as the energy source. The design, concept verification testing and analysis for such a heat exchanger is discussed. The design presented uses a high efficiency compact core to vaporize the oxygen, and in the self-contained unit, attenuates any pressure and flow oscillations which result from unstable boiling in the core. This approach is referred to as the high heat transfer design. An alternative approach which prevents unstable boiling of the oxygen by limiting the heat transfer is referred to as the low heat transfer design and is reported in Pratt & Whitney report FR-19135-2.
Addressed is a class of resonant power processing equipment designed to be used in an integrated high frequency (20 KHz domain), utility power system for large, multi-user spacecraft and other aerospace vehicles. It describes a hardware approach, which has been the basis for parametric and physical data used to justify the selection of high frequency ac as the PMAD baseline for the space station. This paper is part of a larger effort undertaken by NASA and General Dynamics to be sure that all potential space station contractors and other aerospace power system designers understand and can comfortably use this technology, which is now widely used in the commercial sector. In this paper, we will examine control requirements, stability, and operational modes; and their hardware impacts from an integrated system point of view. The current space station PMAD system will provide the overall requirements model to develop an understanding of the performance of this type of system with regard to: (1) regulation; (2) power bus stability and voltage control; (3) source impedance; (4) transient response; (5) power factor effects, and (6) limits and overloads.
A star simulator is described which is capable of simulating a field of three or more stars in response to signals from a computer interface. Star magnitude and spectral output can be manipulated and absolute position measurements can be obtained with the system. The simulator is a single axis system which consists of two assemblies: a light source and filter wheel assembly, and a translation stage assembly. The assemblies are self contained on 60.96 x 60.96 cm breadboards and are connected via three fiber optic cables. A summary of operational modes is provided.
The radiometric calibration accuracy of the Shuttle Imaging Radar (SIR-C) sensor is discussed. The analysis includes the antenna, RF electronics, the digital data handling system, the platform attitude control, attitude determination accuracy, and orbit effects. The radiometric distortion of the image products by the ground processing system used for the image formation is also considered. Since the SIR-C system is a dual-frequency quad-polarized system (i.e., 8 channels), the amplitude and phase error is considered over all possible operating modes and environments for absolute and relative (long-term and short-term) calibration within a channel and across channels.
Results of laboratory experiments with a standard multiwire imaging proportional counter (MWPC), modified to include a parallel grid preamplification stage are presented. Both argon and xenon were used as the primary absorbing gas in conjunction with a number of quench gases. It is found that extremely good energy resolution is obtained from the preamplification stage, almost independent of the ionization potential of the quench gas and over a broad range of gains. Furthermore, regardless of absorbing gas or quench agent, there was no difficulty in transferring a portion of this preamplified charge to the MWPC section for multistep mode operation. Finally, even for overall system gains greater than 10,000, the energy resolution from the second stage is found to be significantly improved, as long as the preamplification stage is operating. These results should have wide application and are of special interest for X-ray astronomy.
A control system utilizing the Multi-mission Modular Spacecraft (MMS) Attitude Control System (ACS) was developed and analyzed for the Space Infrared Telescope Facility (SIRTF) spacecraft. Alternative torquer augmentation schemes were studied to determine viable ACS approaches. A control law was developed to use a dual set of single-axis Control Moment Gyros (CMGs) for two-axis control. Flexible structural models were developed using a high fidelity, flight tested NASTRAN model of the MMS, coupled with a NASTRAN model of the SIRTF telescope. Modal significance criteria were employed to reduce the structural model. Multivariable interactive techniques were used to synthesize the control system (including the structural filters). Control system performance for the SIRTF operational modes (quiescent inertial hold, slewing, nodding, and rastering) was then determined using both single-axis and three-axis simulations). The control system described met performance requirements for all modes but the raster with the use of CMGs. The raster performance was limited by the structural flexibility.
Design features of the attitude control and determination (ACAD) system for the Gamma Ray Observatory (GRO) that will eventually be launched on the Shuttle are described. A tabulation of the ACAD system components is provided and the various standby and normal pointing operational modes of the system are summarized. The system software and sensors will maintain a quaternion model of the GRO attitude on the bases of kinematic equations and inertial data. The software is standardized and has previously been used on the Solar Maximum Mission and Landsat-D. Details of the processing components, redundant electronics for sensor processing, data handling and actuator control are outlined and illustrated with block diagrams. Tests applied to validate the ACAD design are outlined, as are ground support which will be implemented once the GRO is launched.
This paper describes the Orbital Maneuvering Vehicle (OMV) and its intended role as a key element of NASA's space infrastructure. Status, plans, and operational modes are summarized. Typical mission scenarios supporting the servicing of spacecraft and platforms from both the Shuttle and the Space Station are described. Particular emphasis is placed on the orbital mechanics associated with the placement and retrieval of spacecraft and platforms. For example, the optimum placement of a Space Station co-orbiting spacecraft in order to maximize the time interval during which it can be retrieved by a Space Station based OMV is shown as a function of the ballistic coefficient of the spacecraft.
The preliminary design for the Space Telescope Imaging Spectrograph (STIS), which has been selected by NASA for definition study for future flight as a second-generation instrument on the Hubble Space Telescope (HST), is presented. STIS is a two-dimensional spectrograph that will operate from 1050 A to 11,000 A at the limiting HST resolution of 0.05 arcsec FWHM, with spectral resolutions of 100, 1200, 20,000, and 100,000 and a maximum field-of-view of 50 x 50 arcsec. Its basic operating modes include echelle model, long slit mode, slitless spectrograph mode, coronographic spectroscopy, photon time-tagging, and direct imaging. Research objectives are active galactic nuclei, the intergalactic medium, global properties of galaxies, the origin of stellar systems, stelalr spectral variability, and spectrographic mapping of solar system processes.
Results are presented of a comparative analysis of Hettrick Bowyer Type II (HBII) and Wolter-Schwarzschild Type II (WSII) optics for use with Rowland circle spectrometers. The HBII can substitute, with few exceptions, for a WSII in any Rowland circle with little change in spectrometer performance or design. However, the HBII telescope offers several clear advantages over the WSII in these configurations. Because the HBII employs a virtual focus, it requires a much shorter instrument length than a WSII. For example, a 3-m Rowland circle spectrometer, fed by an optimized, f/10, 1-m diameter WSII, has a total instrument length of 6 m. If a HBII is used to feed the identical spectrometer, the entire instrument length can be as little as 3 m. In addition, the improved imaging gained with the larger graze angles of the HBII design results in better resolution in slitless operation modes.