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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

Optical Time Transfer for Bistatic SAR Spacecraft

A spacecraft-to-spacecraft optical time-transfer simulation has been developed as a tool for informing NASA’s Surface Deformation and Change (SDC) mission architecture. The SDC mission will combine radar images from multiple spacecraft to improve understanding of the Earth’s sea-level and landscape changes. Spacecraft must be precisely synchronized in order to create sharp radar images. Simulation of multiple spacecraft time-synchronizing via laser communication can inform technology choices of a mission by providing a picosecond-precision level estimate of clock error. This timing and ranging simulation has been combined with a radar system performance analysis pipeline. The simulated timing errors are used to predict performance of bistatic SAR systems in the presence of oscillator noise and time synchronization in accuracy. This analysis includes both analytic approximation equations from existing literature, and a numerical radar simulation to extract key system performance parameters like phase error and signal-to-noise ratio (SNR)degradation. Precision time-transfer techniques facilitate the accurate synchronization of clocks between any combination of terminals. Most time-transfer technology for comparing two clocks at different terminals use radio frequencies (RF) to measure the time delay between the sending and receiving of signals. Laser technology offers the capability to transmit high data rates with systems that are of smaller size and lower power than comparable RF systems. The clocks on independent spacecraft will have some phase and frequency errors between them that result in clock drift. The two clock models that are included in this bi-directional MATLAB simulation are a cesium-based Chip-Scale Atomic Clock (CSAC) and a rubidium-based Miniature Atomic Clock (MAC). The CSAC has flown as hardware for small satellite missions such as the University of Florida’s CHOMPTT mission. A study of example orbits, including that of NASA NASA-ISRO Synthetic Aperture Radar Mission (NISAR) mission, and lasing rates demonstrate the impact of flight configuration parameters on the synchronization error between two spacecraft. The MATLAB timing simulation uses a Runge-Kutta 4th-order method to propagate spacecraft orbits and computes the light-travel time estimate between them. The simulation outputs the estimated range and estimated clock error based on a user-defined spacecraft cluster configuration. The radar simulation and analytic approximations are applied to evaluate a potential future NASA bistatic SAR constellation architecture. In the proposed architecture, satellites follow each other in the same orbit at 800 km altitude, with a 210 km baseline. We also baseline the CSAC as an ultra stable oscillator, and use NASA’s NISAR for baseline radar system parameters to compute a clock-system introduced phase error of 5.6 degrees without synchronization by frequent time transfer. We build on this base case with a sensitivity analysis of radar performance over a proposed range of constellation and radar system parameters. With this analysis pipeline, we comment on which radar parameters should or should not be changed to minimize synchronization requirements. This analysis technique could be extended or modified to evaluate the timing requirements of other geometries for other future multistatic SAR missions, or other interferometric satellite missions.

Surface Deformation and Change↗
NASA NTRS2008

Return to Mercury: A Comparison of Solar Simulation and Flight Data for the MESSENGER Spacecraft

The MErcury, Surface, Space, ENvironment GEochemistry and Ranging (MESSENGER) spacecraft is a NASA Discovery Mission spacecraft developed and operated by the Johns Hopkins University Applied Physics Laboratory. It was launched on August 3, 2004 and is currently on a course for Mercury orbit insertion in March 2011. To date the mission trajectory has taken the spacecraft to minimum solar distances of 0.332 and 0.313 AU and on January 14, 2008 the first flyby of Mercury in 33 years. From launch through the latest perihelion passage temperature performance data has been collected for the sun facing Digital Sun Sensors (DSS), the sun facing phased array and low gain (omni) antennas, the solar arrays, the sunshade and the two sun facing attitude control 4.4 N thrusters. Prior to launch, extensive solar simulation testing was conducted at the Glenn Research Center, Tank 6 solar simulation facility in Cleveland Ohio. Flight hardware qualification units representing these Sun exposed components were tested in solar environments that represented near mission minimum solar distance as to verify the thermal designs and the material used in fabrication. The paper will review the thermal designs of these components and their thermal performance to date as compared to the solar simulation testing.

Ercol, Carl J.↗
NASA NTRS1972

Advanced ETC/LSS computerized analytical models, CO2 concentration. Volume 1: Summary document

Computer simulations have been prepared for the concepts of C02 concentration which have the potential for maintaining a C02 partial pressure of 3.0 mmHg, or less, in a spacecraft environment. The simulations were performed using the G-189A Generalized Environmental Control computer program. In preparing the simulations, new subroutines to model the principal functional components for each concept were prepared and integrated into the existing program. Sample problems were run to demonstrate the methods of simulation and performance characteristics of the individual concepts. Comparison runs for each concept can be made for parametric values of cabin pressure, crew size, cabin air dry and wet bulb temperatures, and mission duration.

Taylor, B. N.↗
NASA NTRS1965

Solar simulator spectrum evaluation.

Computer method using spectrometer for rapidly measuring spectral distribution of solar simulator and determining effects on thermal balance evaluation of spacecraft

SOLAR SIMULATOR↗
NASA NTRS1995

An Attitude Control Design for the Cassini Spacecraft

The Cassini planetary mission to Saturn and Titan requires accurate pointing and complex maneuvers of the spacecraft, which is comprised of a Saturn orbiter and Titan probe. An attitude control design is described to achieve the many capabilities of Cassini including thrust vector control, precision pointing and stability control for remote sciences, tracking of a probe released to the Titan atmosphere, and radar mapping of the Titan surface. Additional constraints for control design have been imposed by a high percentage of liquid propellant (60% at launch) land lowly damped flexible spacecraft appendages. Computer simulation results of attitude determination, pointing and maneuver controls are provided to demonstrate the design capability.

Cassini↗
NASA NTRS2001

Design of the EO-1 Pulsed Plasma Thruster Attitude Control Experiment

The Pulsed Plasma Thruster (PPT) Experiment on the Earth Observing 1 (EO-1) spacecraft has been designed to demonstrate the capability of a new generation PPT to perform spacecraft attitude control. The PPT is a small, self-contained pulsed electromagnetic Propulsion system capable of delivering high specific impulse (900-1200 s), very small impulse bits (10-1000 micro N-s) at low average power (less than 1 to 100 W). EO-1 has a single PPT that can produce torque in either the positive or negative pitch direction. For the PPT in-flight experiment, the pitch reaction wheel will be replaced by the PPT during nominal EO-1 nadir pointing. A PPT specific proportional-integral-derivative (PID) control algorithm was developed for the experiment. High fidelity simulations of the spacecraft attitude control capability using the PPT were conducted. The simulations, which showed PPT control performance within acceptable mission limits, will be used as the benchmark for on-orbit performance. The flight validation will demonstrate the ability of the PPT to provide precision pointing resolution. response and stability as an attitude control actuator.

Zakrzwski, Charles↗
NASA NTRS1987

Evolution of the solar wind structure in the outer heliosphere

Shocks and interaction regions play very important roles in the evolution of large-scale solar wind structure in the outer heliosphere. This study is based on (1) plasma and magnetic field data observed from Voyager and Pioneer spacecraft, and (2) a quantitative magnetohydrodynamic simulation model. Interaction regions bounded by a forward and a reverse shock begin to form near 1 AU at the leading edges of a large-scale stream. The total pressure in the region is greater than the ambient pressure by a factor of ten or more. Large jumps in pressure remain as a prominant feature of the interplanetary structure even as the jumps in flow speed become less visible in the outer heliosphere. The propagation of the forward and reverse shocks widens the dimension of an interaction region. As a result, two interaction regions belonging to neighboring streams coalesce to form a merged interaction region (MIR). Collision and merging of shocks take place during the coalescence process. Two MIRs can themselves merge again at greater heliocentric distances. Simulation results agree well with spacecraft observations, and they explain major restructuring of the solar wind in the outer heliosphere.

Whang, Y. C.↗
NASA NTRS1994

Progress in navigation filter estimate fusion and its application to spacecraft rendezvous

A new derivation of an algorithm which fuses the outputs of two Kalman filters is presented within the context of previous research in this field. Unlike other works, this derivation clearly shows the combination of estimates to be optimal, minimizing the trace of the fused covariance matrix. The algorithm assumes that the filters use identical models, and are stable and operating optimally with respect to their own local measurements. Evidence is presented which indicates that the error ellipsoid derived from the covariance of the optimally fused estimate is contained within the intersections of the error ellipsoids of the two filters being fused. Modifications which reduce the algorithm's data transmission requirements are also presented, including a scalar gain approximation, a cross-covariance update formula which employs only the two contributing filters' autocovariances, and a form of the algorithm which can be used to reinitialize the two Kalman filters. A sufficient condition for using the optimally fused estimates to periodically reinitialize the Kalman filters in this fashion is presented and proved as a theorem. When these results are applied to an optimal spacecraft rendezvous problem, simulated performance results indicate that the use of optimally fused data leads to significantly improved robustness to initial target vehicle state errors. The following applications of estimate fusion methods to spacecraft rendezvous are also described: state vector differencing, and redundancy management.

Carpenter, J. Russell↗
NASA NTRS1994

New potentials of NIICHIMMASH's thermal vacuum facilities

The potentialities of existing test facilities as to simulating space environment governing factors for spacecraft successful development thermal vacuum testing are analyzed, ways of modernizing existing test facilities and specific proposals on their redesign are considered. The problem of spacecraft (S/C) ground development in simulated external environments, the solution of which started more than 30 years ago, has not lost its urgency today. Stringent requirements on S/C active lifetime under space conditions, module large dimensions, great number of extension elements and complicated mode of their interaction in long mission do not allow S/C designers to abandon ground tests. S/C thermal modes development is a combination of calculations, thermal vacuum tests and actions on improving S/C design and its thermal control system. Traditionally, tests are carried out by stages from component and end unit level verifications to complex tests of modules and S/C as a whole. In our opinion, sufficient correctness of calculated models and experience gained in organizations designing space systems allow to reduce cost and time of autonomous tests. Unfortunately, this is not true for complex (integrated) thermal vacuum tests. More than that, their recent programs include tasks of verifying other (than thermal control system) systems if S/C for operation under space simulated conditions. The outlined circumstances are the main reason for critical review of the potentialities of the existing test base, and of NIICHIMMASH's two large thermal vacuum chambers, first of all. The reasons for and ways of enlargement of these facilities potentially are analyzed and the results attained are described.

Afanassiev, N. A.↗
NASA NTRS2008

SAMPEX Spin Stabilized Mode

The Solar, Anomalous, and Magnetospheric Particle Explorer (SAMPEX), the first of the Small Explorer series of spacecraft, was launched on July 3, 1992 into an 82' inclination orbit with an apogee of 670 km and a perigee of 520 km and a mission lifetime goal of 3 years. After more than 15 years of continuous operation, the reaction wheel began to fail on August 18,2007. With a set of three magnetic torquer bars being the only remaining attitude actuator, the SAMPEX recovery team decided to deviate from its original attitude control system design and put the spacecraft into a spin stabilized mode. The necessary operations had not been used for many years, which posed a challenge. However, on September 25, 2007, the spacecraft was successfully spun up to 1.0 rpm about its pitch axis, which points at the sun. This paper describes the diagnosis of the anomaly, the analysis of flight data, the simulation of the spacecraft dynamics, and the procedures used to recover the spacecraft to spin stabilized mode.

Tsai, Dean C.↗
NASA NTRS2012

Burning Questions in Gravity-Dependent Combustion Science

Building upon a long history of spaceflight and ground based research, NASA's Combustion Science program has accumulated a significant body of accomplishments on the ISS. Historically, NASAs low-gravity combustion research program has sought: to provide a more complete understanding of the fundamental controlling processes in combustion by identifying simpler one-dimensional systems to eliminate the complex interactions between the buoyant flow and the energy feedback to the reaction zone to provide realistic simulation of the fire risk in manned spacecraft and to enable practical simulation of the gravitational environment experienced by reacting systems in future spacecraft. Over the past two decades, low-gravity combustion research has focused primarily on increasing our understanding of fundamental combustion processes (e.g. droplet combustion, soot, flame spread, smoldering, and gas-jet flames). This research program was highly successful and was aided by synergistic programs in Europe and in Japan. Overall improvements were made in our ability to model droplet combustion in spray combustors (e.g. jet engines), predict flame spread, predict soot production, and detect and prevent spacecraft fires. These results provided a unique dataset that supports both an active research discipline and also spacecraft fire safety for current and future spacecraft. These experiments have been conducted using the Combustion Integrated Rack (CIR), the Microgravity Science Glovebox and the Express Rack. In this paper, we provide an overview of the earlier space shuttle experiments, the recent ISS combustion experiments in addition to the studies planned for the future. Experiments in combustion include topics such as droplet combustion, gaseous diffusion flames, solid fuels, premixed flame studies, fire safety, and super critical oxidation processes.

Urban, David↗
NASA NTRS2001

Investigation of the Polytropic Relationship Between Density and Temperature Within Interplanetary Coronal Mass Ejections Using Numerical Simulations

Single-point spacecraft measurements within coronal mass ejections (CMEs) often exhibit a negative correlation between electron density and temperature. At least two opposing interpretations have been suggested for this relationship. If, on one hand, these single spacecraft observations provide direct measures of the polytropic properties of the plasma, then they imply that the polytropic index for the electrons gamma(sub e) is often < 1. Moreover, since the electrons carry the bulk of the pressure (via their significantly higher temperature), this further implies that the dynamics of CME evolution are dominated by an effective polytropic index gamma(sub e)ff < 1. On the other hand, gamma < 1 implies that as the ejecta propagate away from the Sun and expand, they also heat up; a result clearly at odds with in situ observations. In contrast to these CME intervals, many studies have shown that the quiescent solar wind exhibits a positive correlation between electron density and temperature, suggesting that gamma(sub e) > 1. In this study we simulate the evolution of a variety of CME-like disturbances in the solar wind using a one-dimensional, single-fluid model, to address the interpretation of the relationship between electron density and temperature within CMEs at fixed locations in space. Although we strictly impose a polytropic relationship (with gamma = constant) throughout our simulations, we demonstrate that a variety of correlations can exist between density and temperature at fixed points. Furthermore, we demonstrate that the presence of only local uncorrelated random fluctuations in density and temperature can produce a negative correlation. Consequently, we conclude that these single-point observations of negative correlations between electron density and temperature cannot be used to infer the value of gamma(sub e). Instead, we suggest that entropy variations, together with the plasma's tendency to achieve pressure balance with its surroundings, are responsible for the observed profiles.

Riley, Pete↗
NASA NTRS2011

Simulating New Drop Test Vehicles and Test Techniques for the Orion CEV Parachute Assembly System

The Crew Exploration Vehicle Parachute Assembly System (CPAS) project is engaged in a multi-year design and test campaign to qualify a parachute recovery system for human use on the Orion Spacecraft. Test and simulation techniques have evolved concurrently to keep up with the demands of a challenging and complex system. The primary simulations used for preflight predictions and post-test data reconstructions are Decelerator System Simulation (DSS), Decelerator System Simulation Application (DSSA), and Drop Test Vehicle Simulation (DTV-SIM). The goal of this paper is to provide a roadmap to future programs on the test technique challenges and obstacles involved in executing a large-scale, multi-year parachute test program. A focus on flight simulation modeling and correlation to test techniques executed to obtain parachute performance parameters are presented.

Morris, Aaron L.↗
NASA NTRS1991

SeGRAm - A practical and versatile tool for spacecraft trajectory optimization

An implementation of the Sequential Gradient/Restoration Algorithm, SeGRAm, is presented along with selected examples. This spacecraft trajectory optimization and simulation program uses variational calculus to solve problems of spacecraft flying under the influence of one or more gravitational bodies. It produces a series of feasible solutions to problems involving a wide range of vehicles, environments and optimization functions, until an optimal solution is found. The examples included highlight the various capabilities of the program and emphasize in particular its versatility over a wide spectrum of applications from ascent to interplanetary trajectories.

Rishikof, Brian H.↗
NASA NTRS1982

Galileo spin rate control

This paper presents the analysis and design of the spin rate control flight algorithm for the dual-spin Galileo spacecraft. Because of thruster plume contamination constraints, only one poorly located thruster is available for spin rate correction in each spin direction. Hence, firing of any of the spin thrusters has a deleterious effect on spacecraft pointing. A control strategy was developed for achieving the desired spin rate correction while keeping disturbances at acceptable levels. This involved symmetrically distributed multiple thruster burns. Results of extensive tests of the adopted scheme on a computer simulation of the spacecraft are also presented.

Eke, F. O.↗
NASA NTRS2010

Handling Qualities of a Capsule Spacecraft During Atmospheric Entry

A piloted simulation was conducted to study handling qualities for capsule spacecraft entering the Earth s atmosphere. Eight evaluation pilots, including six pilot astronauts, provided Cooper-Harper ratings, workload ratings, and qualitative comments. The simulation began after descending through the atmospheric entry interface point and continued until the drogue parachutes deployed. There were two categories of piloting tasks, both of which required bank angle control. In one task category, the pilot followed a closed-loop bank angle command computed by the backup guidance system to manage g-loads during entry. In the other task category, the pilot used intuitive rules to determine the desired bank angle independently, based on an open-loop schedule of vertical speed, Mach, and total energy specified at several range-to-target gates along the entry trajectory. Pilots were able to accurately track the bank angle guidance commands and steered the capsule toward the recovery site with essentially the same range error as the benchmark autopilot trajectory albeit with substantially higher propellant usage, and the handling qualities for this task were satisfactory. Another key result was that the complex piloting task of atmospheric entry could be performed satisfactorily, even in the presence of large dispersions, by controlling bank angle to follow a simple open-loop schedule.

Bilimoria, Karl D.↗