ALBERT

Description: All Shuttle Solid Rocket Motors (SRM's) exhibit low amplitude longitudinal pressure oscillations during motor burn. Although the oscillations have no known deleterious effect on motor ballistics, the acoustic pressure variations cause thrust oscillations that might affect Shuttle systems or components. The acoustic mode of greatest interest is the first or fundamental mode which, in the SRM, has a nominal frequency of 14-Hz. Oscillations in the SRM are believed to be caused by coupling between large scale vortices and the acoustic modes of the motor chamber. The vortices are thought to be created in the region of the motor segment interfaces and are inherent in the design of the motor. In such a situation the usual approach is to measure the oscillations and assess their impact on any sensitive components through tests and analysis. Questionable components can be altered to survive the vibration environment. As motor firings occur, oscillations are monitored to determine whether there are changes in the nature of the oscillations. Since the first static test, SRM's have been equipped with instrumentation especially designed to acquire chamber pressure oscillation data. Data from the first SRM static tests were used to establish predicted upper bounds for the maximum amplitudes in the latter half of burn. Those bounds have been used as a basis for worst-case simulation scenarios by specialists in structural dynamics at NASA and Rockwell International and to provide a basis for evaluating data from individual motors which were tested subsequent to the original SRM's. This paper updates the upper bounds prediction the High Performance Motors (HPM) by including data from all static tests performed to date including both original SRM's and post Challenger SRM's or Reusable Solid Rocket Motors (RSRM) in which the joint design was changed. All together, this study examines 27 SRM motors, 16 HPM motors and 11 RSRM motors. Predicted upper bounds will be made for both the first and second longitudinal modes. The first mode upper bounds will be compared to the original seven standard rocket motors (STD). The results indicate that, although the upper bounds have increased, they are still within acceptable bounds.

Description: An analysis embodied in a PC computer program is presented, which quantitatively demonstrates how the availability of radiation hard solar cells can help minimize the cost of a global satellite communications system. An important distinction between the currently proposed systems, such as Iridium, Odyssey and Ellipsat, is the number of satellites employed and their operating altitudes. Analysis of the major costs associated with implementing these systems shows that operation at orbital altitudes within the earth's radiation belts (10(exp 3) to 10(exp 4)km) can reduce the total cost of a system by several hundred percent, so long as radiation hard components including solar cells can be used. A detailed evaluation of the predicted performance of photovoltaic arrays using several different planar solar cell technologies is given, including commercially available Si and GaAs/Ge, and InP/Si which is currently under development. Several examples of applying the program are given, which show that the end of life (EOL) power density of different technologies can vary by a factor of ten for certain missions. Therefore, although a relatively radiation-soft technology can usually provide the required EOL power by simply increasing the size of the array, the impact upon the total system budget could be unacceptable, due to increased launch and hardware costs. In aggregate, these factors can account for more than a 10% increase in the total system cost. Since the estimated total costs of proposed global-coverage systems range from $1B to $9B, the availability of radiation-hard solar cells could make a decisive difference in the selection of a particular constellation architecture.

Description: The Office of Advanced Concepts at NASA recently initiated a re-analysis of the 'Solar Power Satellite' (SPS) concept. The ground rule for the analysis was that a project of such size (and more particularly, cost) as the 1980 baseline concept is out of the question in today's world. The study questions are: (1) have technology improvements since the 1980 studies made SPS concepts more feasible?; and (2) are new architectures or concepts for SPS possible which would reduce the cost? These issues were posed to the workshop conducted at the SPRAT Conference, with the intent of soliciting input from experts on space photovoltaic technology.

Description: The atmosphere of Mars has a considerable load of suspended dust. Over time, this dust is deposited out of the atmosphere. The mechanism and the temporal and geographical variation of this deposition are not well characterized. Measurements of settling rates and dust properties are of considerable scientific interest. Atmospheric dust affects the atmospheric solar absorption and thus the heat balance of Mars, as well as serving as nucleation sites for water and CO2 frost. Knowledge of dust properties is of critical interest to design and prediction of the lifetime and power output of solar arrays, and also to design of mechanical mechanisms and radiators. An instrument has been designed and fabricated to measure the dust accumulation during the course of the Mars Pathfinder rover mission. The solar-cell coverglass transmission experiment will measure the change in optical opacity of a transparent coverglass as dust settles on the surface, and a quartz crystal monitor will measure the mass deposited.

Description: As part of the space qualification effort for blue-red reflecting coverslides designed for use with GaAs solar cells, the first long-term (3000 hours) UV testing of unirradiated and 1 MeV electron-irradiated GaAs solar cells, with 4 types of multilayer-coated coverslides to reduce operating temperature, has produced some unexpected results. Important conclusions from this study, which includes two parallel tests, are as follows: (1) All of the GaAs solar cells with multilayer-coated coverslides display UV degradation. The laboratory data, extrapolated to 10 years in orbit, point to a significant loss mechanism from a combination of absorption and a reduction in optical match in such coatings from this portion of the space environment; (2) The effects of contamination in a vacuum system, on the measured degradation in solar-cell short-circuit current during a UV test, depend upon the type of coverslide coatings present on the coverslide surfaces. This has implications for both coated coverslides and optical solar reflectors (OSR's) in space; and (3) Because of the observed trends in this test and uncertainties in the extrapolation of data for multilayer coated coverslides, the use of any multilayer-coated coverslides for extended missions (greater than 1 year) cannot be recommended without prior flight testing.

Description: This paper describes Japanese activities on mainly silicon solar cell research development and applications. The high efficiency thin silicon solar cells and the same kinds of solar cells with integrated bypass function (IBF cells) were developed and qualified for space applications. The most efficient cells (NRS/LBSF cells) showed average 18% at AMO and 28 C conditions. After electron irradiation, NRS/BSF cells showed higher efficiency than NRS/LBSF cells. The IBF cells do not suffer high reverse voltage and can survive from shadowing. The designs and characteristics of these solar cells are presented. In the last section, our future plan for the solar cell calibration is presented.

Description: Several current rocket engine concepts such as the bell-annular tri-propellant engine, and the linear aerospike being proposed for the X-33 require unconventional three dimensional rocket nozzles which must conform to rectangular or sector shaped envelopes to meet integration constraints. These types of nozzles exist outside the current experience database, therefore, the application of efficient design methods for these propulsion concepts is critical to the success of launch vehicle programs. The objective of this work is to optimize several different nozzle configurations, including two- and three-dimensional geometries. Methodology includes coupling computational fluid dynamic (CFD) analysis to genetic algorithms and Taguchi methods as well as implementation of a streamline tracing technique. Results of applications are shown for several geometeries including: three dimensional thruster nozzles with round or super elliptic throats and rectangualar exits, two- and three-dimensional thrusters installed within a bell nozzle, and three dimensional thrusters with round throats and sector shaped exits. Due to the novel designs considered for this study, there is little experience which can be used to guide the effort and limit the design space. With a nearly infinite parameter space to explore, simple parametric design studies cannot possibly search the entire design space within the time frame required to impact the design cycle. For this reason, robust and efficient optimization methods are required to explore and exploit the design space to achieve high performance engine designs. Five case studies which examine the application of various techniques in the engineering environment are presented in this paper.

Description: Rocket propulsion is not ideal when the propellant is not ejected at a unique velocity in an inertial frame. An ideal velocity distribution requires that the exhaust velocity vary linearly with the velocity of the vehicle in an inertial frame. It also requires that the velocity distribution variance as a thermodynamic quantity be minimized. A rocket vehicle with an inert propellant is not optimal, because it does not take advantage of the propellant mass for energy storage. Nor is it logical to provide another energy storage device in order to realize variable exhaust velocity, because it would have to be partly unfilled at the beginning of the mission. Performance is enhanced by pushing on the surrounding because it increases the reaction mass and decreases the reaction jet velocity. This decreases the fraction of the energy taken away by the propellant and increases the share taken by the payload. For an optimal model with the propellant used as fuel, the augmentation realized by pushing on air is greatest for vehicles with a low initial/final mass ratio. For a typical vehicle in the Earth's atmosphere, the augmentation is seen mainly at altitudes below about 80 km. When drag is taken into account, there is a well-defined optimum size for the air intake. Pushing on air has the potential to increase the performance of rockets which pass through the atmosphere. This is apart from benefits derived from "air breathing", or using the oxygen in the atmosphere to reduce the mass of an on-board oxidizer. Because of the potential of these measures, it is vital to model these effects more carefully and explore technology that may realize their advantages.

Description: Pulsed plasma thrusters (PPT's) are a new option for attitude control of a small spacecraft and may result in reduced attitude control system (ACS) mass and cost. The primary purpose of an ACS is to orient the spacecraft configuration to the desired accuracy in inertial space. The ACS functions for which the PPT system will be analyzed include disturbance torque compensation and slewing maneuvers such as sun acquisition for which the small impulse bit and high specific impulse of the PPT offers unique advantages. The NASA Lewis Reserach Center (LeRC) currently has a contracted flight PPT system development program in place with Olin Aerospace and a delivery date of October 1997. The PPT system in this study are based upon the work being done under the NASA LeRC program. Analysis of the use of PPT's for ACS showed that the replacement of the standard momentum wheels and torque rods systems with a PTT system to perform the altitude control maneuvers on a small low Earth orbiting spacecraft reduced the ACS mass by 50 to 75 percent with no increase in required power level over comparable wheel-based systems.

Description: Gravity Probe B, an experiment to test the theory of relativity, will be launched near the turn of the millennium. Due to the precise pointing requirements need to successfully carry out this experiment, the satellite will use sixteen proportionally controlled microthrusters as a main component of the attitude control system. These microthrusters use the helium boil-off from the on-board dewar as propellant. Marshall Space Flight Center, verified the design of the thruster flow path by both computational and experimental methods. The flow performance of the thruster has been adequately characterized. Graphs show specific impulse, thrust coefficient, discharge coefficient, and mass flow rate trends. Value was added to the program through gained confidence in the design of the thruster and through valuation of some design trade-offs.