SEARCH · Search NASA
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.
Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.
Light contact with long-range spacecraft flying between the earth and the moon was considered. The possibility of creating on-board light signaling devices and the problems of construction of light signal receiving devices are studied. It was concluded that signaling using constant burning lamps is most favorable for observation of the spacecraft and transmission of information from the spacecraft.
Summary of spacecraft range safety trajectory analyses for Apollo missions
The structure of spacecraft tracking radar, and range and Doppler measurements, are discussed. Ranging can be used in conjunction with Doppler to determine the total number of charged particles on the radio path. Ranging can also be employed as the primary data type in determining certain kinds of orbits. The basic configuration of the ranging and Doppler system is considered. Inherent limitations on its accuracy with respect to the oscillator, transponder, ground equipment, signal to noise ratio, and general operation are noted.
The Formation Flying Testbed (FFTB) at the National Aeronautics and Space Administration (NASA) Goddard Space Flight Center (GSFC) provides a hardware-in-the-loop test environment for formation navigation and control. The facility is evolving as a modular, hybrid, dynamic simulation facility for end-to-end guidance, navigation and. control (GN&C) design and analysis of formation flying spacecraft. The core capabilities of the FFTB, as a platform for testing critical hardware and software algorithms in-the-loop, have expanded to include S-band Radio Frequency (RF) modems for inter-spacecraft communication and ranging. To enable realistic simulations that require RF ranging sensors for relative navigation, a mechanism is needed to buffer the RF signals exchanged between spacecraft that accurately emulates the dynamic environment through which the RF signals travel, including the effects of medium, moving platforms, and radiated power. The Path Emulator for RF Signals (PERFS), currently under development at NASA GSFC, provides this capability. The function and performance of a prototype device are presented.
Report discusses determination of effective densities of electrically charged particles along paths of microwave signals transmitted to and from spacecraft and use of those densities to calibrate measured signal-propagation times, used to compute distances between ground stations and spacecraft. Oriented toward selection of method of calibration optimal with respect to three criteria: minimization of range error and of root-mean-square uncertainty in range error; minimization of amount, complexity, and/or cost of equipment; and capability of calibrating changes in group velocities of signals on ranging channel.
The Formation Flying Testbed (FFTB) at the National Aeronautics and Space Administration (NASA) Goddard Space Flight Center (GSFC) provides a hardware-in-the-loop test environment for formation navigation and control. The facility is evolving as a modular, hybrid, dynamic simulation facility for end-to-end guidance, navigation and control (GN&C) design and analysis of formation flying spacecraft. The core capabilities of the FFTB, as a platform for testing critical hardware and software algorithms in-the-loop, have expanded to include S-band Radio Frequency (RF) modems for interspacecraft communication and ranging. To enable realistic simulations that require RF ranging sensors for relative navigation, a mechanism is needed to buffer the RF signals exchanged between spacecraft that accurately emulates the dynamic environment through which the RF signals travel, including the effects of the medium, moving platforms, and radiated power. The Path Emulator for Radio Frequency Signals (PERFS), currently under development at NASA GSFC, provides this capability. The function and performance of a prototype device are presented.
One function of the Deep Space Network is validation of the range data that they receive. In this paper we present an automated online sequential range predictor which shows promise of significantly reducing computational and manpower expenditures. The proposed algorithm, a U-D covariance factored Kalman filter, is demonstrated by processing a four-month record of Viking spacecraft data taken enroute to Mars.
Results are presented of analysis of statistical characteristics of information from spacecraft data compression systems, ignoring the nature of channel activity. Analysis of the distribution of the number of active channels and the channel activities was performed using data from scientific experiments conducted on the Zond-1 and Venera-4 spacecraft.
Improved solutions have been obtained for the orbit and equatorial cross-section of Mercury using radar ranging data spanning 22 years. These data have yielded new results on the precession of Mercury's perihelion and better limits on a possible time variation in the gravitational constant G.
Spacecraft range measurements have provided the most accurate tests, to date, of some relativistic gravitational parameters, even though the measurements were made with ranging systems having error budgets of about 10 meters. Technology is now available to allow an improvement of two orders of magnitude in the accuracy of spacecraft ranging. The largest gains in accuracy result from the replacement of unstable analog components with high speed digital circuits having precisely known delays and phase shifts.
Two types of measurements of Mercury's surface topography were obtained by the MESSENGER (MErcury Surface Space ENvironment, GEochemisty and Ranging) spacecraft: laser ranging data from Mercury Laser Altimeter (MLA) [1], and stereo imagery from the Mercury Dual Imaging System (MDIS) camera [e.g., 2, 3]. MLA data provide precise and accurate elevation meaurements, but with sparse spatial sampling except at the highest northern latitudes. Digital terrain models (DTMs) from MDIS have superior resolution but with less vertical accuracy, limited approximately to the pixel resolution of the original images (in the case of [3], 15-75 m). Last year [4], we reported topographic measurements of craters in the D=2.5 to 5 km diameter range from stereo images and suggested that craters on Mercury degrade more quickly than on the Moon (by a factor of up to approximately 10×). However, we listed several alternative explanations for this finding, including the hypothesis that the lower depth/diameter ratios we observe might be a result of the resolution and accuracy of the stereo DTMs. Thus, additional measurements were undertaken using MLA data to examine the morphometry of craters in this diameter range and assess whether the faster crater degradation rates proposed to occur on Mercury is robust.
The relativity test is based on Mariner 9 tracking data acquired by the Deep Space Network. These data consist of two types of measurements: (1) two-way frequency difference (Doppler) measurements that are proportional to the tracking station-spacecraft range rate, and (2) signal round trip time-delay measurements that are proportional to the station-spacecraft range. A data compression scheme has been employed to alleviate accuracy and cost difficulties by (1) using the Doppler data to solve for the spacecraft orbit and to relate spacecraft position to the center of mass of Mars; and (2) combining this result with the station-spacecraft range measurements to obtain measurements of the Earth-Mars distance, called normal points, that contain all of the information pertinent to a detailed knowledge of the Earth-Mars motion.
Binary-coded sequential acquisition spacecraft ranging system operating at weak signals using HF digital logic
Optimum design of tracking system using transponder to determine spacecraft range
In recent years, there has been a tendency toward shrinking the size of spacecraft. New classes of spacecraft called micro-spacecraft have been defined by their mass, power, and size ranges. Spacecraft in the range of 20 to 100 kg represent the class most likely to be utilized by most small sat users in the near future. There are also efforts to develop 10 to 20 kg class spacecraft for use in satellite constellations. More ambitious efforts will be to develop spacecraft less than 10 kg, in which MEMS fabrication technology is required. These new micro-spacecraft will require new micro-propulsion technology. Although micro-propulsion includes electric propulsion approaches, the focus of this proposed program is micro-chemical propulsion which requires the development of microcombustors. As combustors are scaled down, the surface to volume ratio increases. The heat release rate in the combustor scales with volume, while heat loss rate scales with surface area. Consequently, heat loss eventually dominates over heat release when the combustor size becomes smaller, thereby leading to flame quenching. The limitations imposed on chamber length and diameter has an immediate impact on the degree of miniaturization of a micro-combustor. Before micro-combustors can be realized, such a difficulty must be overcome. One viable combustion alternative is to take advantage of surface catalysis. Micro-chemical propulsion for small spacecraft can be used for primary thrust, orbit insertion, trajectory-control, and attitude control. Grouping micro-propulsion devices in arrays will allow their use for larger thrust applications. By using an array composed of hundreds or thousands of micro-thruster units, a particular configuration can be arranged to be best suited for a specific application. Moreover, different thruster sizes would provide for a range of thrust levels (from N s to mN s) within the same array. Several thrusters could be fired simultaneously for thrust levels higher than the basic units, or in a rapid sequence in order to provide gradual but steady low-g acceleration. These arrays of micro-propulsion systems would offer unprecedented flexibility and redundancy for satellite propulsion and reaction control for launch vehicles. A high-pressure bi-propellant micro-rocket engine is already being developed using MEMS technology. High pressure turbopumps and valves are to be incorporated onto the rocket chip . High pressure combustion of methane and O2 in a micro-combustor has been demonstrated without catalysis, but ignition was established with a spark. This combustor has rectangular dimensions of 1.5 mm by 8 mm (hydraulic diameter 3.9 mm) and a length of 4.5 mm and was operated at 1250 kPa with plans to operate it at 12.7 MPa. These high operating pressures enable the combustion process in these devices, but these pressures are not practical for pressure fed satellite propulsion systems. Note that the use of these propellants requires an ignition system and that the use of a spark would impose a size limitation to this micro-propulsion device because the spark unit cannot be shrunk proportionately with the thruster. Results presented in this paper consist of an experimental evaluation of the minimum catalyst temperature for initiating/supporting combustion in sub-millimeter diameter tubes. The tubes are resistively heated and reactive premixed gases are passed through the tubes. Tube temperature and inlet pressure are monitored for an indication of exothermic reactions and composition changes in the gases.
The Return Data Delay technique which requires knowledge of spacecraft range is commonly used for correlating a spacecraft clock against a ground time standard when millisecond accuracy is required. An analysis is presented that allows using the user spacecraft clock calibration system (USCCS) to correlate a spacecraft clock to better than one microsecond accuracy. The basic USCCS algorithm has been simplified and it is shown to result in about one microsecond accuracy without requiring orbital information. By considering the relative motion of the user satellite, the TDRS and the earth station about the center of the earth, a correction of almost two orders of magnitude can be made. Such accuracy is required for scientific investigations that require correlating coincident radiation or particle detection with a remote laboratory.
The current Tracking and Data Relay Satellite configuration consists of nine in-orbit satellites (four first generation, three second generation and two third generation satellites) globally distributed in geosynchronous orbit to provide near continuous data relay service to missions like Hubble Space Telescope and the International Space Station. The 1st generation spacecraft were designed by TRW/Northrop Grumman with their launches of the five spacecraft ranging from 1983 through 1995. The 2nd and 3rd generation spacecraft were designed by Boeing with their launches ranging 2000 - 2002 and 2013 - 2017 respectively. TDRS-3 is now 27 years on orbit, continues to be a capable asset for the TDRS constellation. Lack of need for inclination control combined with large fuel reserves and redundancy on critical elements provides spacecraft that operate well past design life, all of which contributes to expanded TDRS constellation support capabilities. All spacecraft generations have issues. Significant issues will be summarized with the focus on the Boeing related problems. Degradations and failures are continually assessed and provide the foundation for yearly updates to spacecraft reliability models, constellation service projections and deorbit plans (in order to meet NASAs mandate of limiting orbital debris). Even when accounting for degradations and failures, the life expectancy for the Boeing delivered 2nd generation TDRS-8, 9 10 TDRS are anticipated to be 25+ years.