ALBERT

Description: The Department of Energy, Stirling Technology Company (STC), and NASA Glenn Research Center (NASA Glenn) are developing a free-piston Stirling convertor for a high-efficiency Stirling Radioisotope Generator (SRG) for NASA Space Science missions. The SRG is being developed for multimission use, including providing electric power for unmanned Mars rovers and deep space missions. NASA Glenn is conducting an in-house technology project to assist in developing the convertor for space qualification and mission implementation. Recent testing, of 55-We Technology Demonstration Convertors (TDC's) built by STC includes mapping, of a second pair of TDC's, single TDC testing, and TDC electromagnetic interference and electromagnetic compatibility characterization on a nonmagnetic test stand. Launch environment tests of a single TDC without its pressure vessel to better understand the convertor internal structural dynamics and of dual-opposed TDC's with several engineering mounting structures with different natural frequencies have recently been completed. A preliminary life assessment has been completed for the TDC heater head, and creep testing of the IN718 material to be used for the flight convertors is underway. Long-term magnet aging tests are continuing to characterize any potential aging in the strength or demagnetization resistance of the magnets used in the linear alternator (LA). Evaluations are now beginning on key organic materials used in the LA and piston/rod surface coatings. NASA Glenn is also conducting finite element analyses for the LA, in part to look at the demagnetization margin on the permanent magnets. The world's first known integrated test of a dynamic power system with electric propulsion was achieved at NASA Glenn when a Hall-effect thruster was successfully operated with a free-piston Stirling power source. Cleveland State University is developing a multidimensional Stirling computational fluid dynamics code to significantly improve Stirling loss predictions and assist in identifying convertor areas for further improvements. This paper will update the status and results for these efforts.

Description: Ram Burn Observations (RAMBO) is a Department of Defense experiment that observes shuttle Orbital Maneuvering System engine burns for the purpose of improving plume models. On STS-107 the appropriate sensors will observe selected rendezvous and orbit adjust burns.

Description: When building smaller, less expensive spacecraft, there is a need for intelligent fault tolerance vs. increased hardware redundancy. If fault tolerance can be achieved using existing navigation sensors, cost and vehicle complexity can be reduced. A maximum likelihood-based approach to thruster fault detection and identification (FDI) for spacecraft is developed here and applied in simulation to the X-38 space vehicle. The system uses only gyro signals to detect and identify hard, abrupt, single and multiple jet on- and off-failures. Faults are detected within one second and identified within one to five accords,

Description: The prospects of using electrical power beamed from space are coming closer to reality with the continued pursuit and improvements in the supporting space solar research and technology. Space Solar Power (SSP) has been explored off and on for approximately three decades as a viable alternative and clean energy source. Results produced through the more recent Space Solar Power Exploratory Research and Technology (SERT) program involving extensive participation by industry, universities, and government has provided a sound technical basis for believing that technology can be improved to the extent that SSP systems can be built, economically feasible, and successfully deployed in space. Considerable advancements have been made in conceptual designs and supporting technologies including solar power generation, wireless power transmission, power management distribution, thermal management and materials, and the integrated systems engineering assessments. Basic technologies have progressed to the point were the next logical step is to formulate and conduct sophisticated demonstrations involving prototype hardware as final proof of concepts and identify high end technology readiness levels in preparation for full scale SSP systems designs. In addition to continued technical development issues, environmental and safety issues must be addressed and appropriate actions taken to reassure the public and prepare them for the future use of this alternative renewable energy resource. Accomplishing these objectives will allow informed future decisions regarding further SSP and related R&D investments by both NASA management and prospective external partners. In particular, accomplishing these objectives will also guide further definition of SSP and related technology roadmaps including performance objectives, resources and schedules; including 'multi-purpose' applications (terrestrial markets, science, commercial development of space, and other government missions).

Description: Space Solar Power technology offers unique benefits for near-term NASA space science missions, which can mature this technology for other future applications. "Laser-Photo-Voltaic Wireless Power Transmission" (Laser-PV WPT) is a technology that uses a laser to beam power to a photovoltaic receiver, which converts the laser's light into electricity. Future Laser-PV WPT systems may beam power from Earth to satellites or large Space Solar Power satellites may beam power to Earth, perhaps supplementing terrestrial solar photo-voltaic receivers. In a near-term scientific mission to the moon, Laser-PV WPT can enable robotic operations in permanently shadowed lunar polar craters, which may contain ice. Ground-based technology demonstrations are proceeding, to mature the technology for this initial application, in the moon's polar regions.

Description: NASA's In-Space Propulsion (ISP) Program is designed to develop advanced propulsion technologies that can enable or greatly enhance near and mid-term NASA science missions by significantly reducing cost, mass, and/or travel times. These technologies include: Electric Propulsion (Solar and Nuclear Electric) [note: The Nuclear Electric Propulsion work will be transferred to the NSI program in FY03]; Propellantless Propulsion (aerocapture, solar sails, plasma sails, and momentum exchange tethers); Advanced Chemical Propulsion. The ISP approach to identifying and prioritizing these most promising technologies is to use mission analysis and subsequent peer review. These technologies under consideration are mid-Technology Readiness Level (TRL) up to TRL-6 for incorporation into mission planning within three - five years of initiation. In addition, maximum use of open competition is encouraged to seek optimum solutions under ISP. Several NASA Research Announcements (NRAs) have been released asking industry, academia and other organizations to propose propulsion technologies designed to improve our ability to conduct scientific study of the outer planets and beyond. The ISP Program is managed by NASA HQ (Headquarters) and implemented by the Marshall Space Flight Center in Huntsville, Alabama.

Description: The specific heater control requirements for the thermal vacuum and thermal balance testing of the Microwave Anisotropy Probe (MAP) Observatory at the Goddard Space Flight Center (GSFC) in Greenbelt, Maryland are described. The testing was conducted in the 10m wide x 18.3m high Space Environment Simulator (SES) Thermal Vacuum Facility. The MAP thermal testing required accurate quantification of spacecraft and fixture power levels while minimizing heater electrical emissions. The special requirements of the MAP test necessitated construction of five (5) new heater racks.

Description: This viewgraph presentation gives an overview of the turbulent mixing of primary and secondary flow streams in a rocket-based combined cycle (RBCC) engine. A significant RBCC ejector mode database has been generated, detailing single and twin thruster configurations and global and local measurements. On-going analysis and correlation efforts include Marshall Space Flight Center computational fluid dynamics modeling and turbulent shear layer analysis. Potential follow-on activities include detailed measurements of air flow static pressure and velocity profiles, investigations into other thruster spacing configurations, performing a fundamental shear layer mixing study, and demonstrating single-shot Raman measurements.

Description: A LOX/GH2 swirl injector was designed for a 10:1 propellant throttling range. To accomplish this, a dual LOX (liquid oxygen) manifold was used feeding a single common vortex chamber of the swirl element. Hot-fire experiments were conducting for rocket chamber pressures from 80 to 800 psia at a mixture ratio of nominally 6.0 using steady flow, single-point-per-firing cases as well as dynamic throttling conditions. Low frequency (mean) and high frequency (fluctuating) pressure transducer data, flow meter measurements, and Raman spectroscopy images for mixing information were obtained. The injector design, experimental setup, low frequency pressure data, and injector performance analysis will be presented. C efficiency was very high (approximately 100%) at the middle of the throttle-able range with somewhat lower performance at the high and low ends. From the analysis of discreet steady state operating conditions, injector pressure drop was slightly higher than predicted with an inviscid analysis, but otherwise agreed well across the design throttling range. Analysis of the dynamic throttling data indicates that the injector may experience transient conditions that effect pressure drop and performance when compared to steady state results.

Description: The ratio of doubly to singly charged ions was measured in the plumes of a 30 cm and of a 40 cm ion thruster. The measured ratio was correlated with observed erosion rates and thruster operating conditions. The measured and calculated erosion rates paralleled variation in the j(sup ++)/j(sup +) ratio and indicated that the erosion was dominated by Xe III. Simple models of cathode potential surfaces which were developed in support of this work were in agreement with this conclusion and provided a predictive capability of the erosion given the ratio of doubly to singly charged ion currents.

Description: NASA Glenn Research Center is developing a 5/10-kW ion engine for a broad range of mission applications. Simultaneously, a 5-kW breadboard poster processing unit is being designed and fabricated. The design includes a beam supply consisting of four 1.1 kW power modules connected in parallel, equally sharing the output current. A novel phase-shifted/pulse-width-modulated dual full-bridge topology was chosen for its soft-switching characteristics. The proposed modular approach allows scalability to higher powers as well as the possibility of implementing an N+1 redundant beam supply. Efficiencies in excess of 96% were measured during testing of a breadboard beam power module. A specific mass of 3.0 kg/kW is expected for a flight PRO. This represents a 50% reduction from the state of the art NSTAR power processor.

Description: Data were obtained from a 30 cm xenon ion thruster in which the accelerator grid was translated in the radial plane. The thruster was operated at three different throttle power levels, and the accelerator grid was incrementally translated in the X, Y, and azimuthal directions. Plume data was obtained downstream from the thruster using a Faraday probe mounted to a positioning system. Successive probe sweeps revealed variations in the plume direction. Thruster perveance, electron backstreaming limit, accelerator current, and plume deflection angle were taken at each power level, and for each accelerator grid position. Results showed that the thruster plume could easily be deflected up to six degrees without a prohibitive increase in accelerator impingement current. Results were similar in both X and Y direction.

Description: With recent success of the NSTAR ion thruster on Deep Space 1, there is continued interest in long term, high propellant throughput thrusters to perform energetic missions. This requires flight qualified thrusters that can operate for long periods at high beam density, without degradation in performance resulting from sputter induced grid erosion. Carbon-based materials have shown nearly an order of magnitude improvement in sputter erosion resistance over molybdenum. NASA Glenn Research Center (GRC) has been active over the past several years pursuing carbon-based grid development. In 1995, NASA GRC sponsored work performed by the Jet Propulsion Laboratory to fabricate carbon/carbon composite grids using a machined panel approach. In 1999, a contract was initiated with a commercial vendor to produce carbon/carbon composite grids using a chemical vapor infiltration process. In 2001, NASA GRC purchased pyrolytic carbon grids from a commercial vendor. More recently, a multi-year contract was initiated with North Carolina A&T to develop carbon/carbon composite grids using a resin injection process. The following paper gives a brief overview of these four programs.

Description: A series of tests were conducted to measure the combustion performance of the Fastrac engine thrust chamber. During mainstage, the thrust chamber exhibited no large-amplitude chamber pressure oscillations that could be identified as low-frequency combustion instability or 'chug'. However, during start-up and shutdown, the thrust chamber very briefly exhibited large-amplitude chamber pressure oscillations that were identified as chug. These instabilities during start-up and shutdown were regarded as benign due to their brevity. Linear models of the thrust chamber and the propellant feed systems were formulated for both the thrust chamber component tests and the flight engine tests. These linear models determined the frequency and decay rate of chamber pressure oscillations given the design and operating conditions of the thrust chamber and feed system. The frequency of chamber pressure oscillations determined from the model closely matched the frequency of low-amplitude, low-frequency chamber pressure oscillations exhibited in some of the later thrust chamber mainstage tests. The decay rate of the chamber pressure oscillations determined from the models indicated that these low-frequency oscillations were stable. Likewise, the decay rate, determined from the model of the flight engine tests indicated that the low-frequency chamber pressure oscillations would be stable.

Description: The NASA Glenn Research Center is developing hydrogen based combined cycle propulsion technology for a single-stage-to-orbit launch vehicle application under a project called GTX. Rocket Based Combined Cycle (RBCC) propulsion systems incorporate one or more rocket engines into an airbreathing flow path to increase specific impulse as compared to an all rocket-powered vehicle. In support of this effort, an RBCC direct-connect test capability was established at the Engine Components Research Laboratory to investigate low speed, ejector ramjet, and initial ramjet operations and performance. The facility and test article enables the evaluation of two candidate low speed operating schemes; the simultaneous mixing and combustion (SMC) and independent ramjet stream (IRS). The SMC operating scheme is based on the fuel rich operations of the rocket where performance depends upon mixing between the rocket plume and airstream. In contrast, the IRS scheme fuels the airstream separately and uses the rocket plume to ignite the fuel-air mixture. This paper describes the test hardware and facility upgrades installed to support the RBCC tests. It also defines and discusses low speed technical challenges being addressed by the experiments. Finally, preliminary test results, including rocket risk mitigating tests, unfueled airflow tests, and the integrated system hot fire test will be presented.

Description: Electron backstreaming due to accelerator grid hole enlargement has been identified as a failure mechanism that will limit ion thruster service lifetime. Over extended periods of time as accelerator grid apertures enlarge due to erosion, ion thrusters are required to operate at increasingly higher accelerator grid voltages in order to prevent electron backstreaming. These higher voltages give rise to higher grid erosion rates, which in turn accelerates aperture enlargement. Presented here is an approach to mitigate the backflow of electrons into the engine by using a transverse magnetic field.

Description: Economy of scale is inherent in the microwave power transmission aperture/spot-size trade-off, resulting in a requirement for large space systems in the existing design concepts. Unfortunately, this large size means that the initial investment required before the first return, and the price of amortization of this initial investment, is a daunting (and perhaps insurmountable) barrier to economic viability. As the growth of ground-based solar power applications will fund the development of the PV technology required for space solar power and will also create the demand for space solar power by manufacturing a ready-made market, space power systems must be designed with an understanding that ground-based solar technologies will be implemented as a precursor to space-based solar. for low initial cost, (3) operation in synergy with ground solar systems, and (4) power production profile tailored to peak rates. A key to simplicity of design is to maximize the integration of the system components. Microwave, millimeter-wave, and laser systems are analyzed. A new solar power satellite design concept with no sun-tracking and no moving parts is proposed to reduce the required cost to initial operational capability.

Description: The availability of abundant, affordable power where needed is a key to the future exploration and development of space as well as future sources of clean terrestrial power. One innovative approach to providing such power is the use of wireless power transmission (WPT). There are at least two possible WPT methods that appear feasible; microwave and laser. Microwave concepts have been generated, analyzed and demonstrated. Technologies required to provide an end-to-end system have been identified and roadmaps generated to guide technology development requirements. Recently, laser W T approaches have gained an increased interest. These approaches appear to be very promising and will possibly solve some of the major challenges that exist with the microwave option. Therefore, emphasis is currently being placed on the laser WPT activity. This paper will discuss the technology requirements, technology roadmaps and technology flight experiments demonstrations required to lead toward a pilot plant demonstration. Concepts will be discussed along with the modeling techniques that are used in developing them. Feasibility will be addressed along with the technology needs, issues and capabilities for particular concepts. Flight experiments and demonstrations will be identified that will pave the road from demonstrations to pilot plants and beyond.

Description: The objective of the FRC (Field Reversed Configuration) Acceleration Space Thruster (FAST) Experiment is to investigate the use of a repetitive FRC source as a thruster, specifically for an NEP (nuclear electric propulsion) system. The Field Reversed Configuration is a plasmoid with a closed poloidal field line structure, and has been extensively studied as a fusion reactor core. An FRC thruster works by repetitively producing FRCs and accelerating them to high velocity. An FRC thruster should be capable of I(sub sp)'s in the range of 5,000 - 25,000 seconds and efficiencies in the range of 60 - 80 %. In addition, they can have thrust densities as high as 10(exp 6) N/m2, and as they are inductively formed, they do not suffer from electrode erosion. The jet-power should be scalable from the low to the high power regime. The FAST experiment consists of a theta-pinch formation chamber, followed by an acceleration stage. Initially, we will produce and accelerate single FRCs. The initial focus of the experiment will be on the ionization, formation and acceleration of a single plasmoid, so as to determine the likely efficiency and I(sub sp). Subsequently, we will modify the device for repetitive burst-mode operation (5-10 shots). A variety of diagnostics are or will be available for this work, including a HeNe interferometer, high-speed cameras, and a Thomson-scattering system. The status of the experiment will be described.

Description: The Mini-Magnetospheric Plasma Propulsion (M2P2), originally proposed by Winglee et al. [2000], is based on the two-fluid plasma model and requires a 15-km frontal standoff distance (or 20-km cross-sectional diameter) in order for the magnetic bubble to absorb sufficient momentum from the SW to accelerate a spacecraft to the unprecedented speeds of 50-80 km/s after an acceleration period of about three months. Winglee et al. [2000] derived the above size requirement based on an extrapolation of their simulated results in which a system much smaller than a M2P2 was used (p. 21,074 of their study). We submit, however, that a fluid model has no validity for such a small scale size-even in the region near the plasma source! It is assumed in the MHD fluid model, normally applied to the magnetosphere, that the characteristic scale-size is much greater than the Larmor radius and ion skin depth of the SW. In the case of the M2P2, however, the size of the magnetic bubble is actually less than or, at best, comparable to, the scale of these characteristic parameters and, therefore, a kinetic approach, which addresses the smallscale physical mechanisms involved, must be used. A fully three-dimensional version of the hybrid code is used in our M2P2 (Plasma Sails) studies was originally developed by Delamere et al. [1999]. The M2P2 plasma sail is an excellent application for this hybrid code. The primary advantage of this code is the seamless interface between fluid and kinetic descriptions of the ion populations. A kinetic description is not necessary for the dense inner regions of the magnetic bubble and tremendous computational savings can be realized by treating this dense, magnetized ion population with the fluid description. It is essential, however, that the outer bubble regions be treated kinetically as well as the SW protons. Comparison of full size M2P2 simulation based on 3D MHD and kinetic models show that kinetic treatment introduces much more asymmetry to the considering problem and the possibility of kinetic instabilities development.

Description: In this paper we present a comparison of optimization approaches to the minimum fuel rendezvous problem. Both indirect and direct methods are compared for a variety of test cases. The indirect approach is based on primer vector theory. The direct approaches are implemented numerically and include Sequential Quadratic Programming (SQP), Quasi-Newton, Simplex, Genetic Algorithms, and Simulated Annealing. Each method is applied to a variety of test cases including, circular to circular coplanar orbits, LEO to GEO, and orbit phasing in highly elliptic orbits. We also compare different constrained optimization routines on complex orbit rendezvous problems with complicated, highly nonlinear constraints.

Description: This paper examines the approach taken to building a low-cost, modular spacecraft bus that can be used to support a variety of technology experiments in different space environments. It describes the techniques used and design drivers considered to ensure experiment independence from as yet selected host spacecraft. It describes the technology experiment carriers that will support NASA's Living With a Star Space Environment Testbed space missions. NASA has initiated the Living With a Star (LWS) Program to develop a better scientific understanding to address the aspects of the connected Sun-Earth system that affect life and society. A principal goal of the program is to bridge the gap between science, engineering, and user application communities. The Space Environment Testbed (SET) Project is one element of LWS. The Project will enable future science, operational, and commercial objectives in space and atmospheric environments by improving engineering approaches to the accommodation and/or mitigation of the effects of solar variability on technological systems. The SET Project is highly budget constrained and must seek to take advantage of as yet undetermined partnering opportunities for access to space. SET will conduct technology validation experiments hosted on available flight opportunities. The SET Testbeds will be developed in a manner that minimizes the requirements for accommodation, and will be flown as flight opportunities become available. To access the widest range of flight opportunities, two key development requirements are to maintain flexibility with respect to accommodation constraints and to have the capability to respond quickly to flight opportunities. Experiments, already developed to the technology readiness level of needing flight validation in the variable Sun-Earth environment, will be selected on the basis of the need for the subject technology, readiness for flight, need for flight resources and particular orbit. Experiments will be accumulated by the Project and manifested for specific flight opportunities as they become available. The SET Carrier is designed to present a standard set of interfaces to SET technology experiments and to be modular and flexible enough to interface to a variety of possible host spacecraft. The Carrier will have core components and mission unique components. Once the core carrier elements have been developed, only the mission unique components need to be defined and developed for any particular mission. This approach will minimize the mission specific cost and development schedule for a given flight opportunity. The standard set of interfaces provided by SET to experiments allows them to be developed independent of the particulars of a host spacecraft. The Carrier will provide the power, communication, and the necessary monitoring features to operate experiments. The Carrier will also provide all of the mechanical assemblies and harnesses required to adapt experiments to a particular host. Experiments may be hosted locally with the Carrier or remotely on the host spacecraft. The Carrier design will allow a single Carrier to support a variable number of experiments and will include features that support the ability to incrementally add experiments without disturbing the core architecture.

Description: Nuclear fusion appears to be a most promising concept for producing extremely high specific impulse rocket engines. One particular fusion concept which seems to be very well suited for fusion propulsion applications is the gasdynamic mirror (GDM). An experimental GDM device has been constructed at the NASA Marshall Space Flight Center to provide an initial assessment of the feasibility of this type of propulsion system. A systems shakedown of the device is currently underway with initial experiments slated to occur in early 2002. The device has been constructed so as to allow a considerable degree of flexibility in its configuration thus permitting the experiment to easily grow over time without necessitating a great deal of additional fabrication. This flexibility is due in large part to the modular nature of the machine wherein additional modules may be added as needed to meet varying experimental objectives. Figure 1 shows the current configuration of the Gasdynamic Mirror Experiment, and Table 1 describes the main features of the device.

Description: Flight times and deliverable masses for electric and fusion propulsion systems are difficult to approximate. Numerical integration is required for these continuous thrust systems. Many scientists are not equipped with the tools and expertise to conduct interplanetary and interstellar trajectory analysis for their concepts. Several charts plotting the results of well-known trajectory simulation codes were developed and are contained in this paper. These charts illustrate the dependence of time of flight and payload ratio on jet power, initial mass, specific impulse and specific power. These charts are intended to be a tool by which people in the propulsion community can explore the possibilities of their propulsion system concepts. Trajectories were simulated using the tools VARITOP and IPOST. VARITOP is a well known trajectory optimization code that involves numerical integration based on calculus of variations. IPOST has several methods of trajectory simulation; the one used in this paper is Cowell's method for full integration of the equations of motion. An analytical method derived in the companion paper was also evaluated. The accuracy of this method is discussed in the paper.

Description: Most fusion propulsion concepts that have been investigated in the past employ some form of inertial or magnetic confinement separately, and are encumbered by the need for advanced drivers (e.g. laser) or steady-state magnetic confinement systems (e.g. superconductors) that have historically resulted in large, massive spacecraft designs. Here we present a comparatively new approach, Magnetized Target Fusion (MTF), which offers a nearer-term avenue for realizing the tremendous performance benefits of fusion propulsion. MTF attempts to combine the favorable attributes of both inertially and magnetically confined fusion to achieve both efficient and low-cost compressional plasma heating and energy confinement. The key advantage of MTF is its less demanding requirements for driver energy and power processing. Additional features include: 1) very low system masses and volumes, 2) relatively low waste heat, 3) substantial utilization of energy from product neutrons, 4) efficient, low peak-power drivers based on existing pulsed power technology, 5) very high Isp , specific power and thrust, and 6) relatively affordable R&D pathways. MTF overcomes many of the problems associated with traditional fusion techniques, thus making it particularly attractive for space applications. Isp greater than 50,000 seconds and specific powers greater than 20 kilowatts/kilogram appear feasible using relatively near-term pulse power and plasma gun technology.

Description: The current research effort at NASA Marshall Space Flight Center (MSFC) in MTF is directed towards exploring the critical physics issues of potential embodiments of MTF for propulsion, especially standoff drivers involving plasma liners for MTF. There are several possible approaches for forming plasma liners. One approach consists of using a spherical array of plasma jets to form a spherical plasma shell imploding towards the center of a magnetized plasma, a compact toroid. Current experimental plan and status to explore the physics of forming a 2-D plasma liner (shell) by merging plasma jets are described. A first-generation coaxial plasma guns (Mark-1) to launch the required plasma jets have been built and tested. Plasma jets have been launched reproducibly with a low jitter, and velocities in excess of 50 km/s for the leading edge of the plasma jet. Some further refinements are being explored for the plasma gun, Successful completion of these single-gun tests will be followed by an experimental exploration of the problems of launching a multiple number of these jets simultaneously to form a cylindrical plasma liner.

Description: Next-generation, regeneratively cooled rocket engines require materials that can meet high temperatures while resisting the corrosive oxidation-reduction reaction of combustion known as blanching, the main cause of engine failure. A project was initiated at NASA-Marshal Space Flight Center (MSFC) to combine three existing technologies to build and demonstrate an advanced liquid rocket engine combustion chamber that would provide a 100 mission life. Technology developed in microgravity research to build cartridges for space furnaces was utilized to vacuum plasma spray (VPS) a functional gradient coating on the hot wall of the combustion liner as one continuous operation, eliminating any bondline between the coating and the liner. The coating was NiCrAlY, developed previously as durable protective coatings on space shuttle high pressure fuel turbopump (HPFTP) turbine blades. A thermal model showed that 0.03 in. NiCrAlY applied to the hot wall of the combustion liner would reduce the hot wall temperature 200 F, a 20% reduction, for longer life. Cu-8Cr-4Nb alloy, which was developed by NASA-Glenn Research Center (GRC), and which possesses excellent high temperature strength, creep resistance, and low cycle fatigue behavior combined with exceptional thermal stability, was utilized as the liner material in place of NARloy-Z. The Cu-8Cr-4Nb material exhibits better mechanical properties at 650 C (1200 F) than NARloy-Z does at 538 C (1000 F). VPS formed Cu-8Cr-4Nb combustion chamber liners with a protective NiCrAlY functional gradient coating have been hot fire tested, successfully demonstrating a durable coating for the first time. Hot fire tests along with tensile and low cycle fatigue properties of the VPS formed combustion chamber liners and witness panel specimens are discussed.

Description: Electromagnetic thrusters typically use electric and magnetic fields to accelerate and exhaust plasma through interactions with the charged particles in the plasma. The energy required to create the plasma, i.e. ionization energy, is potential energy between the electron and ion. This potential energy is typically lost since it is not recovered as the plasma is exhausted and is known as frozen flow loss. If the frozen flow energy is a small fraction of the total plasma energy, this frozen flow loss may be negligible. However, if the frozen flow energy is a major fraction of the total plasma energy, this loss can severely reduce the energy efficiency of the thruster. Recovery and utilization of this frozen flow energy can improve the energy efficiency of a thruster during low specific impulse operating regimes when the ionization energy is a large fraction of the total plasma energy. This paper quantifies the recovery of the frozen flow energy, i.e. recombination energy, via the process of surface recombination for helium. To accomplish this task the momentum flux and heat flux of the plasma flow were measured and compared to calculated values from electrostatic probe data. This information was used to deduce the contribution of recombination energy to the total heat flux on a flat plate as well as to characterize the plasma conditions. Helium propellant was investigated initially due to its high ionization potential and hence available recombination energy.

Description: As the space applications become more complex and timing constraints on control actions are more stringent, the task of integrating and testing NASA's real-time systems (such as X-38 Crew Return Vehicle, and certain International Space Station autonomous systems) has become a great challenge. A testing environment where can preserve consistent temporal behaviors as in the target execution must be established for system-level verification and software quality assurance. Our goal is to develop an analysis suite for validation and verification of real-time systems that are used to perform human- in-the-loop control operations during safety-critical missions. The suite will be able to carry out quantitative approaches of coverage diagnostic and temporal behavior evaluation in order to measure test coverage, to optimize test utilization, and to verify timing correctness.

Description: There are currently three dominant TSTO class architectures. These are Series Burn (SB), Parallel Burn with crossfeed (PBw/cf), and Parallel Burn without crossfeed (PBncf). The goal of this study was to determine what factors uniquely affect PBncf architectures, how each of these factors interact, and to determine from a performance perspective whether a PBncf vehicle could be competitive with a PBw/cf or SB vehicle using equivalent technology and assumptions. In all cases, performance was evaluated on a relative basis for a fixed payload and mission by comparing gross and dry vehicle masses of a closed vehicle. Propellant combinations studied were LOX: LH2 propelled orbiter and booster (HH) and LOX: Kerosene booster with LOX: LH2 orbiter (KH). The study conclusions were: 1) a PBncf orbiter should be throttled as deeply as possible after launch until the staging point. 2) a detailed structural model is essential to accurate architecture analysis and evaluation. 3) a PBncf TSTO architecture is feasible for systems that stage at mach 7. 3a) HH architectures can achieve a mass growth relative to PBw/cf of 〈 20%. 3b) KH architectures can achieve a mass growth relative to Series Burn of 〈 20%. 4) center of gravity (CG) control will be a major issue for a PBncf vehicle, due to the low orbiter specific thrust to weight ratio and to the position of the orbiter required to align the nozzle heights at liftoff. 5 ) thrust to weight ratios of 1.3 at liftoff and between 1.0 and 0.9 when staging at mach 7 appear to be close to ideal for PBncf vehicles. 6) performance for all vehicles studied is better when staged at mach 7 instead of mach 5. The study showed that a Series Burn architecture has the lowest gross mass for HH cases, and has the lowest dry mass for KH cases. The potential disadvantages of SB are the required use of an air-start for the orbiter engines and potential CG control issues. A Parallel Burn with crossfeed architecture solves both these problems, but the mechanics of a large bipropellant crossfeed system pose significant technical difficulties. Parallel Burn without crossfeed vehicles start both booster and orbiter engines on the ground and thus avoid both the risk of orbiter air-start and the complexity of a crossfeed system. The drawback is that the orbiter must use 20% to 35% of its propellant before reaching the staging point. This induces a weight penalty in the orbiter in order to carry additional propellant, which causes a further weight penalty in the booster to achieve the same staging point. One way to reduce the orbiter propellant consumption during the first stage is to throttle down the orbiter engines as much as possible. Another possibility is to use smaller or fewer engines. Throttling the orbiter engines soon after liftoff minimizes CG control problems due to a low orbiter liftoff thrust, but may result in an unnecessarily high orbiter thrust after staging. Reducing the number or size of engines size may cause CG control problems and drift at launch. The study suggested possible methods to maximize performance of PBncf vehicle architectures in order to meet mission design requirements.

Description: The need for fusion propulsion for interplanetary flights is discussed. For a propulsion system, there are three important system attributes: (1) The absolute amount of energy available, (2) the propellant exhaust velocity, and (3) the jet power per unit mass of the propulsion system (specific power). For efficient and affordable human exploration of the solar system, propellant exhaust velocity in excess of 100 km/s and specific power in excess of 10 kW/kg are required. Chemical combustion obviously cannot meet the requirement in propellant exhaust velocity. Nuclear fission processes typically result in producing energy in the form of heat that needs to be manipulated at temperatures limited by materials to about 2,800 K. Using the fission energy to heat a low atomic weight propellant produces propellant velocity of the order of 10 kinds. Alternatively the fission energy can be converted into electricity that is used to accelerate particles to high exhaust velocity. However, the necessary power conversion and conditioning equipment greatly increases the mass of the propulsion system. Fundamental considerations in waste heat rejection and power conditioning in a fission electric propulsion system place a limit on its jet specific power to the order of about 0.2 kW/kg. If fusion can be developed for propulsion, it appears to have the best of all worlds - it can provide the largest absolute amount of energy, the propellant exhaust velocity (〉 100 km/s), and the high specific jet power (〉 10 kW/kg). An intermediate step towards fusion propulsion might be a bimodal system in which a fission reactor is used to provide some of the energy to drive a fusion propulsion unit. There are similarities as well as differences between applying fusion to propulsion and to terrestrial electrical power generation. The similarities are the underlying plasma and fusion physics, the enabling component technologies, the computational and the diagnostics capabilities. These physics and engineering capabilities have been demonstrated for a fusion reactor gain (Q) of the order of unity (TFTR: 0.25, JET: 0.65, JT-60: Q(sub eq) approx. 1.25). These technological advances made it compelling for considering fusion for propulsion.

Description: A radioisotope power and cooling system is designed to provide electrical power for a probe operating on the surface of Venus. Most foreseeable electronics devices and sensors cannot operate at the 450 C ambient surface temperature of Venus. Because the mission duration is substantially long and the use of thermal mass to maintain an operable temperature range is likely impractical, some type of active refrigeration may be required to keep electronic components at a temperature below ambient. The fundamental cooling parameters are the cold sink temperature, the hot sink temperature, and the amount of heat to be removed. In this instance, it is anticipated that electronics would have a nominal operating temperature of 300 C. Due to the highly thermal convective nature of the high-density (90 bar CO2) atmosphere, the hot sink temperature was assumed to be 50 C, which provided a 500 C temperature of the cooler's heat rejecter to the ambient atmosphere. The majority of the heat load on the cooler is from the high temperature ambient surface environment on Venus, with a small contribution of heat generation from electronics and sensors. Both thermoelectric (RTG) and dynamic power conversion systems were analyzed, based on use of a standard isotope (General-purpose heat source, or GPHS) brick. For the radioisotope Stirling power converter configuration designed, the Sage model predicts a thermodynamic power output capacity of 478.1 watts, which slightly exceeds the required 469.1 watts. The hot sink temperature is 1200 C, and the cold sink temperature is 500 C. The required heat input is 1740 watts. This gives a thermodynamic efficiency of 27.48 %. It is estimated that the mechanical efficiency of the power converter design is on the order of 85 %, based on experimental measurements taken from 500-watt power class, laboratory-tested Stirling engines. The overall efficiency is calculated to be 23.36 %. The mass of the power converter is estimated at approximately 21.6 kg. Additional information is included in the original extended abstract.

Description: NASA's Stennis Space Center (SSC) and Glenn Research Center (GRC) participate in the development of technologies for propulsion testing and propulsion applications in air and space transportation. Future transportation systems and the test facilities needed to develop and sustain them are becoming increasingly complex. Sensor technology is a fundamental pillar that makes possible development of complex systems that must operate in automatic mode (closed loop systems), or even in assisted-autonomous mode (highly self-sufficient systems such as planetary exploration spacecraft). Hence, a great deal of effort is dedicated to develop new sensors and related technologies to be used in research facilities, test facilities, and in vehicles and equipment. This paper describes sensor technologies being developed and in use at SSC and GRC, including new technologies in integrated health management involving sensors, components, processes, and vehicles.

Description: This paper addresses the progress of technology development of a laser ignition system at NASA Marshall Space Flight Center (MSFC). Initial hot-fire tests in a small-scale rocket chamber at MSFC have demonstrated the DPLIS concept having two main advantages over existing laser ignition concepts. First, the DPLIS can be potentially optimized its laser pulse format to maximize the initial plasma volume, the plasma lifetime, as well as the flame kernel growth rate. Characterization studies of the laser pulse format are now underway with experiments of igniting gaseous hydrogen/air in a Hencken burner. Once ignition is achieved, the flame is open to the atmosphere. This open environment allows easy access for diagnostics of the ignition phenomenon. The quick turn-around time of conducting experiments on this burner make it more amenable for conducting a large number of experiments for statistical analysis of the sensitivity of the test parameters. The results from these experiments will help optimize the laser format for future testing in an H2/O2 subscale rocket chamber.

Description: This viewgraph representation provides an overview of research which develops a quasi one dimensional chemistry computational fluid dynamics code to study the effect of nozzle design on the performance of pulse detonation rocket engines (PDREs). Topics considered include: PDREs vs. steady-state rocket engines (SSREs), PDRE cycles, numerical models of idealized PDRE performance, thrust determination of PDRE, specific geometries, and nozzle design and geometry.

Description: The National Aeronautics and Space Administration's (NASA) Marshall Space Flight Center (MSFC) is concentrating research into the utilization of photonic materials for spacecraft propulsion. Spacecraft propulsion, using photonic materials, will be achieved using a solar sail. A sail operates on the principle that photons, originating from the sun, impart pressure and provide a source of spacecraft propulsion. The pressure can be increased, by a factor of two if the sun-facing surface is perfectly reflective. Solar sails are generally composed of a highly reflective metallic front layer, a thin polymeric substrate, and occasionally a highly emissive back surface. The Space Environmental Effects Team at MSFC is actively characterizing candidate solar sail materials to evaluate the thermo-optical and mechanical properties after exposure to a simulated Geosynchronous Transfer Orbit (GTO) radiation environment. The technique of radiation dose verses material depth profiling was used to determine the orbital equivalent exposure doses. The solar sail exposure procedures and results of the material characterization will be discussed.

Description: This viewgraph presentation provides information on the structure and activities of the panels of the Joint Army Navy NASA Air Force (JANNAF) Rocket Nozzle Technology Subcommittee. The panels profiled are the Processing Science and Materials Panel, the Nozzle Design, Test, and Evaluation Panel, the Nozzle Analysis and Modeling Panel, and the Nozzle Control Systems Panel. The presentation also lists meetings, workshops, and publications in which the subcommittee participated during the reporting period.

Description: It is very important to accurately predict the gas pressure, gas and solid temperature, as well as the amount of O-ring erosion inside the space shuttle Reusable Solid Rocket Motor (RSRM) joints in the event of a leak path. The scenarios considered are typically hot combustion gas rapid pressurization events of small volumes through narrow and restricted flow paths. The ideal method for this prediction is a transient three-dimensional computational fluid dynamics (CFD) simulation with a computational domain including both combustion gas and surrounding solid regions. However, this has not yet been demonstrated to be economical for this application due to the enormous amount of CPU time and memory resulting from the relatively long fill time as well as the large pressure and temperature rising rate. Consequently, all CFD applications in RSRM joints so far are steady-state simulations with solid regions being excluded from the computational domain by assuming either a constant wall temperature or no heat transfer between the hot combustion gas and cool solid walls.

Description: Recent advances in crystalline solar cell technology are reviewed. Dual-junction and triple-junction solar cells are presently available from several U. S. vendors. Commercially available triple-junction cells consisting of GaInP, GaAs, and Ge layers can produce up to 27% conversion efficiency in production lots. Technology status and performance figures of merit for currently available photovoltaic arrays are discussed. Three specific NASA mission applications are discussed in detail: Mars surface applications, high temperature solar cell applications, and integrated microelectronic power supplies for nanosatellites.

Description: Extending ion engine technology beyond the current state-of-the art primary interplanetary electric propulsion system, the 2.3-kW NASA Solar Electric Propulsion Technology and Applications Readiness (NSTAR) system, will require thrusters with improved propellant throughput and total impulse capability. Many of the design choices that culminated in the NSTAR thrusters must be revisited, and their application to next generation ion engine technology must be evaluated. The concept of derating, which was successfully employed in NSTAR, has been applied to the 40 cm NASA Evolutionary Xenon Thruster (NEXT) currently under development at NASA Glenn Research Center (GRC). At 5-kW, NEXT operates with the same average beam current density as NSTAR, and at 10-kW, the peak beam current density is only ten percent greater than NSTAR. The result is that similar Ion optics technology is expected to yield comparable lifetime. Thick-accelerator- grid ion optics are also being tested to realize additional lifetime benefits. A 40-A discharge cathode is being developed for NEXT based on scaling the NSTAR design. Nevertheless, the experiences of the NSTAR ground tests and the thruster on the Deep Space One spacecraft indicate that the discharge cathode wear must be studied experimentally and theoretically to ensure that it meets the lifetime requirements. Although NEXT is in its infancy, investigations have already begun to examine possible modifications to engine design for even higher-power and higher-specific impulse engines. Ion optics using alternate materials such as titanium, graphite, or carbon-carbon composite are currently being investigated due to their low sputter yields at high voltage. To avoid the difficulties encountered using electrodes at high-currents, the use of a microwave-based ion thruster is under investigation for potential high-power ion thruster systems requiring long lifetimes. Additionally, alternative propellants are being considered for applications requiring high-specific impulse (〉〉 5000 s) and extremely long-life (〉〉 15,000 hr). Testing requirements make condensable propellants attractive for high-power engines. Although the NSTAR ion engine demonstrated the flight maturity of ion thruster technology, many challenges remain for the development of thrusters with improved propellant throughput and power handling capabilities.

Description: A computer model is under continuing development at NASA Glenn Research Center that enables first-order assessments of space power technology. The model, an evolution of NASA Glenn's Array Design Assessment Model (ADAM), is an Excel workbook that consists of numerous spreadsheets containing power technology performance data and sizing algorithms. Underlying the model is a number of databases that contain default values for various power generation, energy storage and power management and distribution component parameters. These databases are actively maintained by a team of systems analysts so that they contain state-of-art data as well as the most recent technology performance projections. Sizing of the power subsystems can be accomplished either by using an assumed mass specific power (W/kg) or energy (Wh/kg) or by a bottoms-up calculation that accounts for individual component performance and masses. The power generation, energy storage and power management and distribution subsystems are sized for given mission requirements for a baseline case and up to three alternatives. This allows four different power systems to be sized and compared using consistent assumptions and sizing algorithms. The component sizing models contained in the workbook are modular so that they can be easily maintained and updated. All significant input values have default values loaded from the databases that can be over-written by the user. The default data and sizing algorithms for each of the power subsystems are described in some detail. The user interface and workbook navigational features are also discussed. Finally, an example study case that illustrates the model's capability is presented.

Description: The photovoltaic (PV) module on the International Space Station (ISS) has been operating since November 2000 and supporting electric power demands of the ISS and its crew of three. The PV module contains photovoltaic arrays that convert solar energy to electrical power and an integrated equipment assembly (IEA) that houses electrical hardware and batteries for electric power regulation and storage. Each PV module contains two independent power channels for fault tolerance. Each power channel contains three batteries in parallel to meet its performance requirements and for fault tolerance. Each battery consists of 76 Ni-Hydrogen (Ni-H2) cells in series. These 76 cells are contained in two orbital replaceable units (ORU) that are connected in series. On-orbit data are monitored and trended to ensure that all hardware is operating normally. Review of on-orbit data showed that while five batteries are operating very well, one is showing signs of mismatched ORUs. The cell pressure in the two ORUs differs by an amount that exceeds the recommended range. The reason for this abnormal behavior may be that the two ORUs have different use history. An assessment was performed and it was determined that capacity of this battery would be limited by the lower pressure ORU. Steps are being taken to reduce this pressure differential before battery capacity drops to the point of affecting its ability to meet performance requirements. As a first step, a battery reinitialization procedure was developed to reduce this pressure differential. The procedure was successfully carried out on-orbit and the pressure differential was reduced to the recommended range. This paper describes the battery performance and the consequences of mismatched ORUs that make a battery. The paper also describes the reinitialization procedure, how it was performed on orbit, and battery performance after the reinitialization. On-orbit data monitoring and trending is an ongoing activity and it will continue as ISS assembly progresses.

Description: Fission technology can enable rapid, affordable access to any point in the solar system. If fission propulsion systems are to be developed to their full potential; however, near-term customers must be identified and initial fission systems successfully developed, launched, and operated. Studies conducted in fiscal year 2001 (IISTP, 2001) show that fission electric propulsion (FEP) systems with a specific mass at or below 50 kg/kWjet could enhance or enable numerous robotic outer solar system missions of interest. At the required specific mass, it is possible to develop safe, affordable systems that meet mission requirements. To help select the system design to pursue, eight evaluation criteria were identified: system integration, safety, reliability, testability, specific mass, cost, schedule, and programmatic risk. A top-level comparison of four potential concepts was performed: a Testable, Passive, Redundant Reactor (TPRR), a Testable Multi-Cell In-Core Thermionic Reactor (TMCT), a Direct Gas Cooled Reactor (DGCR), and a Pumped Liquid Metal Reactor.(PLMR). Development of any of the four systems appears feasible. However, for power levels up to at least 500 kWt (enabling electric power levels of 125-175 kWe, given 25-35% power conversion efficiency) the TPRR has advantages related to several criteria and is competitive with respect to all. Hardware-based research and development has further increased confidence in the TPRR approach. Successful development and utilization of a "Phase I" fission electric propulsion system will enable advanced Phase 2 and Phase 3 systems capable of providing rapid, affordable access to any point in the solar system.

Description: NASA is considering upgrading the Space Shuttle by adding a fifth segment (FSB) to the current four-segment solid rocket booster. Course materials cover design and engineering issues related to the Reusable Solid Rocket Motor (RSRM) raised by the addition of a fifth segment to the rocket booster. Topics cover include: four segment vs. five segment booster, abort modes, FSB grain design, erosive burning, enhanced propellant burn rate, FSB erosive burning model development and hardware configuration.

Description: This paper describes the requirements, design, integration, test, performance, and lessons learned of NASA's Microwave Anisotropy Probe (MAP) propulsion subsystem. MAP was launched on a Delta-II launch vehicle from NASA's Kennedy Space Center on June 30, 2001. Due to instrument thermal stability requirements, the Earth-Sun L2 Lagrange point was selected for the mission orbit. The L2 trajectory incorporated phasing loops and a lunar gravity assist. The propulsion subsystem's requirements are to manage momentum, perform maneuvers during the phasing loops to set up the lunar swingby, and perform stationkeeping at L2 for 2 years. MAP's propulsion subsystem uses 8 thrusters which are located and oriented to provide attitude control and momentum management about all axes, and delta-V in any direction without exposing the instrument to the sun. The propellant tank holds 72 kg of hydrazine, which is expelled by unregulated blowdown pressurization. Thermal management is complex because no heater cycling is allowed at L2. Several technical challenges presented themselves during I and T, such as in-situ weld repairs and in-situ bending of thruster tubes to accommodate late changes in the observatory CG. On-orbit performance has been nominal, and all phasing loop, mid-course correction, and stationkeeping maneuvers have been successfully performed to date.

Description: Experimental investigations are ongoing to study the force imparted to materials when subjected to laser ablation. When a laser pulse of sufficient energy density impacts a material, a small amount of the material is ablated. A torsion balance is used to measure the momentum produced by the ablation process. The balance consists of a thin metal wire with a rotating pendulum suspended in the middle. The wire is fixed at both ends. Recently, multi-layered material systems were investigated. These multi-layered materials were composed of a transparent front surface and opaque sub surface. The laser pulse penetrates the transparent outer surface with minimum photon loss and vaporizes the underlying opaque layer.

Description: This conference presentation reports on the progress on NASA's On-Orbit Propulsion System Project which aims to support the development of second generation reusable launch vehicles (RLV) through advanced research and development and risk reduction activities. Topics covered include: project goals, project accomplishments, risk reduction activities, thruster design and development initiatives, and Aerojet LOX/Ethanol engine development and testing.

Description: This viewgraph presentation provides information on NASA's attempts to develop an air-breathing propulsion in an effort to make future space transportation safer, more reliable and significantly less expensive than today's missions. Spacecraft powered by air-breathing rocket engines would be completely reusable, able to take off and land at airport runways and ready to fly again within days. A radical new engine project is called the Integrated System Tests of an Air-breathing Rocket, or ISTAR.

Description: The exploration of our solar system will require spacecraft with much greater capability than spacecraft which have been launched in the past. This is particularly true for exploration of the outer planets. Outer planet exploration requires shorter trip times, increased payload mass, and ability to orbit or land on outer planets. Increased capability requires better propulsion systems, including increased specific impulse. Chemical propulsion systems are not capable of delivering the performance required for exploration of the solar system. Future propulsion systems will be applied to a wide variety of missions with a diverse set of mission requirements. Many candidate propulsion technologies have been proposed but NASA resources do not permit development of a] of them. Therefore, we need to rationally select a few propulsion technologies for advancement, for application to future space missions. An effort was initiated to select and prioritize candidate propulsion technologies for development investment. The results of the study identified Aerocapture, 5 - 10 KW Solar Electric Ion, and Nuclear Electric Propulsion as high priority technologies. Solar Sails, 100 Kw Solar Electric Hall Thrusters, Electric Propulsion, and Advanced Chemical were identified as medium priority technologies. Plasma sails, momentum exchange tethers, and low density solar sails were identified as high risk/high payoff technologies.

Description: The motivation behind an advanced technology program to develop intelligent power management and distribution (PMAD) systems is described. The program concentrates on developing digital control and distributed processing algorithms for PMAD components and systems to improve their size, weight, efficiency, and reliability. Specific areas of research in developing intelligent DC-DC converters and distributed switchgear are described. Results from recent development efforts are presented along with expected future benefits to the overall PMAD system performance.

Description: A computer simulation of a flywheel energy storage single axis attitude control system is described. The simulation models hardware which will be experimentally tested in the future. This hardware consists of two counter rotating flywheels mounted to an air table. The air table allows one axis of rotational motion. An inertia DC bus coordinator is set forth that allows the two control problems, bus regulation and attitude control, to be separated. Simulation results are presented with a previously derived flywheel bus regulator and a simple PID attitude controller.

Description: Accurate International Space Station (ISS) power prediction requires the quantification of solar array shadowing. Prior papers have discussed the NASA Glenn Research Center (GRC) ISS power system tool SPACE (System Power Analysis for Capability Evaluation) and its integrated shadowing algorithms. On-orbit telemetry has become available that permits the correlation of theoretical shadowing predictions with actual data. This paper documents the comparison of a shadowing metric (total solar array current) as derived from SPACE predictions and on-orbit flight telemetry data for representative significant shadowing cases. Images from flight video recordings and the SPACE computer program graphical output are used to illustrate the comparison. The accuracy of the SPACE shadowing capability is demonstrated for the cases examined.

Description: Results are presented from a parametric assessment of the applicability and spacecraft-level impacts of very lightweight thin-film solar arrays with relatively large deployed areas for representative space missions. The most and least attractive features of thin-film solar arrays are briefly discussed. A calculation is then presented illustrating that from a solar array alone mass perspective, larger arrays with less efficient but lighter thin-film solar cells can weigh less than smaller arrays with more efficient but heavier crystalline cells. However, a spacecraft-level systems assessment must take into account the additional mass associated with solar array deployed area: the propellant needed to desaturate the momentum accumulated from area-related disturbance torques and to perform aerodynamic drag makeup reboost. The results for such an assessment are presented for a representative low Earth orbit (LEO) mission, as a function of altitude and mission life, and a geostationary Earth orbit (GEO) mission. Discussion of the results includes a list of specific mission types most likely to benefit from using thin-film arrays. The presentation concludes with a list of issues to be addressed prior to use of thin-film solar arrays in space and the observation that with their unique characteristics, very lightweight arrays using efficient, thin film cells on flexible substrates may become the best array option for a subset of Earth orbiting and deep space missions.

Description: Successful development of space fission systems requires an extensive program of affordable and realistic testing. In addition to tests related to design/development of the fission system, realistic testing of the actual flight unit must also be performed. If the system is designed to operate within established radiation damage and fuel burn up limits while simultaneously being designed to allow close simulation of heat from fission using resistance heaters, high confidence in fission system performance and lifetime can be attained through non-nuclear testing. Through demonstration of systems concepts (designed by DOE National Laboratories) in relevant environments, this philosophy has been demonstrated through hardware testing in the High Power Propulsion Thermal Simulator (HPPTS). The HPPTS is designed to enable very realistic non-nuclear testing of space fission systems. Ongoing research at the HPPTS is geared towards facilitating research, development, system integration, and system utilization via cooperative efforts with DOE labs, industry, universities, and other NASA centers. Through hardware based design and testing, the HPPTS investigates High Power Electric Propulsion (HPEP) component, subsystem, and integrated system design and performance.

Description: Bipropellant propulsion systems currently represent the largest bus subsystem for many missions. These missions range from low Earth orbit satellite to geosynchronous communications and planetary exploration. The payoff of high performance bipropellant systems is illustrated by the fact that Aerojet Redmond has qualified a commercial NTO/MMH engine based on the high Isp technology recently delivered by this program. They are now qualifying a NTO/hydrazine version of this engine. The advanced rhenium thrust chambers recently provided by this program have raised the performance of earth storable propellants from 315 sec to 328 sec of specific impulse. The recently introduced rhenium technology is the first new technology introduced to satellite propulsion in 30 years. Typically, the lead time required to develop and qualify new chemical thruster technology is not compatible with program development schedules. These technology development programs must be supported by a long term, Base R&T Program, if the technology s to be matured. This technology program then addresses the need for high performance, storable, on-board chemical propulsion for planetary rendezvous and descent/ascent. The primary NASA customer for this technology is Space Science, which identifies this need for such programs as Mars Surface Return, Titan Explorer, Neptune Orbiter, and Europa Lander. High performance (390 sec) chemical propulsion is estimated to add 105% payload to the Mars Sample Return mission or alternatively reduce the launch mass by 33%. In many cases, the use of existing (flight heritage) propellant technology is accommodated by reducing mission objectives and/or increasing enroute travel times sacrificing the science value per unit cost of the program. Therefore, a high performance storable thruster utilizing fluorinated oxidizers with hydrazine is being developed.

Description: The current cost to launch payloads to low earth orbit (LEO) is approximately loo00 U.S. dollars ($) per pound ($22000 per kilogram). This high cost limits our ability to pursue space science and hinders the development of new markets and a productive space enterprise. This enterprise includes NASA's space launch needs and those of industry, universities, the military, and other U.S. government agencies. NASA's Advanced Space Transportation Program (ASTP) proposes a vision of the future where space travel is as routine as in today's commercial air transportation systems. Dramatically lower launch costs will be required to make this vision a reality. In order to provide more affordable access to space, NASA has established new goals in its Aeronautics and Space Transportation plan. These goals target a reduction in the cost of launching payloads to LEO to $lo00 per pound ($2200 per kilogram) by 2007 and to $100'~ per pound by 2025 while increasing safety by orders of magnitude. Several programs within NASA are addressing innovative propulsion systems that offer potential for reducing launch costs. Various air-breathing propulsion systems currently are being investigated under these programs. The NASA Aerospace Propulsion and Power Base Research and Technology Program supports long-term fundamental research and is managed at GLenn Research Center. Currently funded areas relevant to space transportation include hybrid hyperspeed propulsion (HHP) and pulse detonation engine (PDE) research. The HHP Program currently is addressing rocket-based combined cycle and turbine-based combined cycle systems. The PDE research program has the goal of demonstrating the feasibility of PDE-based hybrid-cycle and combined cycle propulsion systems that meet NASA's aviation and access-to-space goals. The ASTP also is part of the Base Research and Technology Program and is managed at the Marshall Space Flight Center. As technologies developed under the Aerospace Propulsion and Power Base Research and Technology Program mature, they are incorporated into ASTP. One example of this is rocket-based combined cycle systems that are being considered as part of ASTP. The NASA Ultra Efficient Engine Technology (UEET) Program has the goal of developing propulsion system component technology that is relevant to a wide range of vehicle missions. In addition to subsonic and supersonic speed regimes, it includes the hypersonic speed regime. More specifically, component technologies for turbine-based combined cycle engines are being developed as part of UEET.

Description: A thrust stand has been built to measure the force-noise produced by electrostatic micro-Newton (muN) thrusters. The LISA mission's Disturbance Reduction System (DRS) requires thrusters that are capable of producing continuous thrust levels between 1-100 muN with a resolution of 0.1 muN. The stationary force-noise produced by these thrusters must not exceed 0.1 muN/dHz in the measurement bandwidth 10(exp -4) to 1 Hz. The LISA Thrust Stand (LTS) is a torsion-balance type thrust stand designed to meet the following requirements: stationary force-noise measurements from l0( -4) to 1 Hz with 0.1 muN/dHz sensitivity, absolute thrust measurements from 1-100 muN with better than 0.1 muN resolution, and dynamic thruster response from to 10 Hz. The LTS employs a unique vertical configuration, autocollimator for angular position measurements, and electrostatic actuators that are used for dynamic pendulum control and null-mode measurements. Force-noise levels are measured indirectly by characterizing the thrust stand as a spring-mass system. The LTS was initially designed to test the indium FEEP thruster developed by the Austrian Research Center in Seibersdorf (ARCS), but can be modified for testing other thrusters of this type.

Description: Recent improvements in energy densities of batteries open the possibility of using electric rather that hydraulic actuators in space vehicle systems. However, the systems usually require short-duration, high-power pulses. This power profile requires the battery system to be sized to meet the power requirements rather than stored energy requirements, often resulting in a large and inefficient energy storage system. Similar transient power applications have used a combination of two or more disparate energy storage technologies. For instance, placing a capacitor and a battery side-by-side combines the high energy density of a battery with the high power performance of a capacitor and thus can create a lighter and more compact system. A parametric study was performed to identify favorable scenarios for using capacitors. System designs were then carried out using equivalent circuit models developed for five commercial electrochemical capacitor products. Capacitors were sized to satisfy peak power levels and consequently "leveled" the power requirement of the battery, which can then be sized to meet system energy requirements. Simulation results clearly differentiate the performance offered by available capacitor products for the space vehicle applications.

Description: STS-104, launched July 2001, marked the first flight of a single Block 2 Space Shuttle Main Engine (SSME). This new configuration of the SSME is the culmination of well over a decade of gradual engine system upgrades. The launch and mission were a success. However, in the process of post-launch data analysis a Main Propulsion System (MPS) anomaly was noted and tied directly to the shutdown of the Block 2 SSME. An investigation into this anomaly was organized across NASA facilities and across the various hardware component contractors. This paper is a very brief summary of the eventual understanding of the root causes of the anomaly and the process whereby an appropriate mitigation action was proposed. An analytical model of the High Pressure Fuel Pump (HPFP) and the low pressure fuel system of the SSME is presented to facilitate the presentation of this summary. The proposed mitigation action is discussed and, with the launch of STS-108 in November 2001, successfully demonstrated under flight conditions.

Description: A computational investigation has been completed to assess the capability of the NASA Tetrahedral Unstructured Software System (TetrUSS) for simulation of exhaust nozzle flows. Three configurations were chosen for this study: (1) a fluidic jet effects model, (2) an isolated nacelle with a supersonic cruise nozzle, and (3) a fluidic pitchthrust- vectoring nozzle. These configurations were chosen because existing data provided a means for measuring the ability of the TetrUSS flow solver USM3D for simulating complex nozzle flows. Fluidic jet effects model simulations were compared with structured-grid CFD (computational fluid dynamics) data at Mach numbers from 0.3 to 1.2 at nozzle pressure ratios up to 7.2. Simulations of an isolated nacelle with a supersonic cruise nozzle were compared with wind tunnel experimental data and structured-grid CFD data at Mach numbers of 0.9 and 1.2, with a nozzle pressure ratio of 5. Fluidic pitch-thrust-vectoring nozzle simulations were compared with static experimental data and structured-grid CFD data at static freestream conditions and nozzle pressure ratios from 3 to 10. A fluidic injection case was computed with the third configuration at a nozzle pressure ratio of 4.6 and a secondary pressure ratio of 0.7. Results indicate that USM3D with the S-A turbulence model provides accurate exhaust nozzle simulations at on-design conditions, but does not predict internal shock location at overexpanded conditions or pressure recovery along a boattail at transonic conditions.

Description: The annular combustor geometry of a combined-cycle engine has been analyzed with three-dimensional computational fluid dynamics. Both subsonic combustion and supersonic combustion flowfields have been simulated. The subsonic combustion analysis was executed in conjunction with a direct-connect test rig. Two cold-flow and one hot-flow results are presented. The simulations compare favorably with the test data for the two cold flow calculations; the hot-flow data was not yet available. The hot-flow simulation Indicates that the conventional ejector-ramjet cycle would not provide adequate mixing at the conditions tested. The supersonic combustion ramjet flowfield was simulated with frozen chemistry model. A five-parameter test matrix was specified, according to statistical design-of-experiments theory. Twenty-seven separate simulations were used to assemble surrogate models for combustor mixing efficiency and total pressure recovery. Scramjet injector design parameters (injector angle, location, and fuel split) as well as mission variables (total fuel mass flow and freestream Mach number) were included in the analysis. A promising injector design has been identified that provides good mixing characteristics with low total pressure losses. The surrogate models can be used to develop performance maps of different injector designs. Several complex three-way variable interactions appear within the dataset that are not adequately resolved with the current statistical analysis.

Description: The first U.S. power module on International Space Station (ISS) was activated in December 2000. Comprised of solar arrays, nickel-hydrogen (NiH2) batteries, and a direct current power management and distribution (PMAD) system, the electric power system (EPS) supplies power to housekeeping and user electrical loads. Modeling EPS performance is needed for several reasons, but primarily to assess near-term planned and off-nominal operations and because the EPS configuration changes over the life of the ISS. The System Power Analysis for Capability Evaluation (SPACE) computer code is used to assess the ISS EPS performance. This paper describes the process of validating the SPACE EPS model via ISS on-orbit telemetry. To accomplish this goal, telemetry was first used to correct assumptions and component models in SPACE. Then on-orbit data was directly input to SPACE to facilitate comparing model predictions to telemetry. It will be shown that SPACE accurately predicts on-orbit component and system performance. For example, battery state-of-charge was predicted to within 0.6 percentage points over a 0 to 100 percent scale and solar array current was predicted to within a root mean square (RMS) error of 5.1 Amps out of a typical maximum of 220 Amps. First, SPACE model predictions are compared to telemetry for the ISS EPS components: solar arrays, NiH2 batteries, and the PMAD system. Second, SPACE predictions for the overall performance of the ISS EPS are compared to telemetry and again demonstrate model accuracy.

Description: The Microwave Anisotropy Probe is a follow-on to the Differential Microwave Radiometer instrument on the Cosmic Background Explorer. Sixteen months before launch, it was discovered that from the time of the critical design review, configuration changes had resulted in a significant migration of the spacecraft's center of mass. As a result, the spacecraft no longer had a viable backup control mode in the event of a failure of the negative pitch axis thruster. Potential solutions to this problem were identified, such as adding thruster plume shields to redirect thruster torque, adding mass to, or removing it from, the spacecraft, adding an additional thruster, moving thrusters, bending thrusters (either nozzles or propellant tubing), or accepting the loss of redundancy for the thruster. The impacts of each solution, including effects on the mass, cost, and fuel budgets, as well as schedule, were considered, and it was decided to bend the thruster propellant tubing of the two roll control thrusters, allowing that pair to be used for back-up control in the negative pitch axis. This paper discusses the problem and the potential solutions, and documents the hardware and software changes that needed to be made to implement the chosen solution. Flight data is presented to show the propulsion system on-orbit performance.

Description: Nuclear electric propulsion (NEP) offers significant benefits to missions for outer planet exploration. Reaching outer planet destinations, especially beyond Jupiter, is a struggle against time and distance. For relatively near missions, such as a Europa lander, conventional chemical propulsion and NEP offer similar performance and capabilities. For challenging missions such as a Pluto orbiter, neither chemical nor solar electric propulsion are capable while NEP offers acceptable performance. Three missions are compared in this paper: Europa lander, Pluto orbiter, and Titan sample return, illustrating how performance of conventional and advanced propulsion systems vary with increasing difficulty. The paper presents parametric trajectory performance data for NEP. Preliminary mass/performance estimates are provided for a Europa lander and a Titan sample return system, to derive net payloads for NEP. The NEP system delivers payloads and ascent/descent spacecraft to orbit around the target body, and for sample return, delivers the sample carrier system from Titan orbit to an Earth transfer trajectory. A representative scientific payload 500 kg was assumed, typical for a robotic mission. The resulting NEP systems are 100-kWe class, with specific impulse from 6000 to 9000 seconds.

Description: The results of performance tests with two 40 cm ion optics sets are presented and compared to those of 30 cm ion optics with similar aperture geometries. The 40 cm ion optics utilized both NSTAR and TAG (Thick-Accelerator-Grid) aperture geometries. All 40 cm ion optics tests were conducted on a NEXT (NASA's Evolutionary Xenon Thruster) laboratory model ion engine. Ion optics performance tests were conducted over a beam current range of 1.20 to 3.52 A and an engine input power range of 1.1 to 6.9 kW. Measured ion optics' performance parameters included near-field radial beam current density profiles, impingement-limited total voltages, electron backstreaming limits, screen grid ion transparencies, beam divergence angles, and start-up transients. Impingement-limited total voltages for 40 cm ion optics with the NSTAR aperture geometry were 60 to 90 V lower than those with the TAG aperture geometry. This difference was speculated to be due to an incomplete burn-in of the TAG ion optics. Electron backstreaming limits for the 40 cm ion optics with the TAG aperture geometry were 8 to 19 V higher than those with the NSTAR aperture geometry due to the thicker accelerator grid of the TAG geometry. Because the NEXT ion engine provided beam flatness parameters that were 40 to 63 percent higher than those of the NSTAR ion engine, the 40 cm ion optics outperformed the 30 cm ion optics.

Description: During ground tests of electric microthrusters and space tests of electrodynamic tethers the electron emitters must successfully operate at environmental pressures possibly as high as 1x10(exp -4) Pa. High partial pressures of oxygen, nitrogen, and water vapor are expected in such environments. A textured platinum surface was used in this work for field emission cathode assessments because platinum does not form oxide films at low temperatures. Although a reproducible cathode conditioning process did not evolve from this work, some short term tests for periods of 1 to 4 hours showed no degradation of emission current at an electric field of 8 V/mm and background pressures of about 1x10(exp -6) Pa. Increases of background pressure by air flow to about 3x10(exp -4) Pa yield a hostile environment for the textured platinum field emission cathode.

Description: Space Flight hardware requires high power conversion efficiencies due to limited power availability and weight penalties of cooling systems. The International Space Station (ISS) Electric Power System (EPS) DC-DC Converter Unit (DDCU) power converter is no exception. This paper explores the design methods and tradeoffs that were utilized to accomplish high efficiency in the DDCU. An isolating DC to DC converter was selected for the ISS power system because of requirements for separate primary and secondary grounds and for a well-regulated secondary output voltage derived from a widely varying input voltage. A flyback-current-fed push-pull topology or improved Weinberg circuit was chosen for this converter because of its potential for high efficiency and reliability. To enhance efficiency, a non-dissipative snubber circuit for the very-low-Rds-on Field Effect Transistors (FETs) was utilized, redistributing the energy that could be wasted during the switching cycle of the power FETs. A unique, low-impedance connection system was utilized to improve contact resistance over a bolted connection. For improved consistency in performance and to lower internal wiring inductance and losses a planar bus system is employed. All of these choices contributed to the design of a 6.25 KW regulated dc to dc converter that is 95 percent efficient. The methodology used in the design of this DC to DC Converter Unit may be directly applicable to other systems that require a conservative approach to efficient power conversion and distribution.

Description: With the first United States (U.S.) photovoltaic array (PVA) activated on International Space Station (ISS) in December 2000, on-orbit data can now be compared to analytical predictions. Due to ISS operational constraints, it is not always possible to point the front side of the arrays at the Sun. Thus, in many cases, sunlight directly illuminates the backside of the PVA as well as albedo illumination on either the front or the back. During this time, appreciable power is produced since the solar cells are mounted on a thin, solar transparent substrate. It is important to present accurate predictions for both front and backside power generation for mission planning, certification of flight readiness for a given mission, and on-orbit mission support. To provide a more detailed assessment of the ISS power production capability, the authors developed a PVA electrical performance model applicable to generalized bifacial illumination conditions. On-orbit PVA performance data were also collected and analyzed. This paper describes the ISS PVA performance model, and the methods used to reduce orbital performance data. Analyses were performed using SPACE. a NASA-GRC developed computer code for the ISS program office. Results showed a excellent comparison of on-orbit performance data and analytical results.

Description: With advances in carbon composite material, magnetic bearings, microprocessors, and high-speed power switching devices, work has begun on a space qualifiable Energy Momentum Wheel (EMW). An EMW is a device that can be used on a satellite to store energy, like a chemical battery, and manage angular momentum, like a reaction wheel. These combined functions are achieved by the simultaneous and balanced operation of two or more energy storage flywheels. An energy storage flywheel typically consists of a carbon composite rotor driven by a brushless DC motor/generator. Each rotor has a relatively large angular moment of inertia and is suspended on magnetic bearings to minimize energy loss. The use of flywheel batteries on spacecraft will increase system efficiencies (mass and power), while reducing design-production time and life-cycle cost. This paper will present a discussion of flywheel battery design considerations and a simulation of spacecraft system performance utilizing four flywheel batteries to combine energy storage and momentum management for a typical LEO satellite. A proposed set of control laws and an engineering animation will also be presented. Once flight qualified and demonstrated, space flywheel batteries may alter the architecture of most medium and high-powered spacecraft.

Description: A flywheel system and its operator interface are described. Measurements of magnetic bearing negative stiffness are performed. Two digital magnetic bearing control algorithms (PD and estimator based) are defined and their implementations are described. Tuning of each controller is discussed. Comparison of the two controllers' stability, damping noise, and operating current are described. Results describing the superiority of the estimator-based controller are presented and discussed.

Description: An analysis of the acoustical and flowfield environment for the scaled 1-pound-force (lbf) thrust tabletop motor was performed. The jet characterization is based on computational fluid dynamics (CFD) in conjunction with Kirchhoff surface integral formulation and compared with correlations developed for measured rocket noise and a pressure fluctuation scaling (PFS) method. Comparisons are made for the overall sound pressure levels (OASPL's) and spectral dependence of sound pressure level (SPL).

Description: The NASA Hubble Space Telescope (HST) was designed to be deployed and later serviced for maintenance and upgrades, as required, by the space shuttle fleet, with a Goodyear mission life for the batteries. HST was deployed 380 miles above the Earth, from Space Shuttle Discovery, on April 25, 1990. Four servicing missions, (SM1, SM2, SM3A, AND SM3B) have been performed. Astronauts have replaced or modified optics, solar arrays, a power control unit, and various science packages. A fifth Servicing Mission, SM4 scheduled for early 2004, is planned to replace the batteries for the first time. The HST is powered by solar array wings and nickel hydrogen (NiH2) Duracell batteries, which are grouped into two parallel battery modules of three parallel batteries each. With a design life of 7 years at launch, these batteries have surpassed 12 years in orbit, which gives HST the highest number of charge/discharge cycles of any NiH2 battery currently in low earth orbit (LEO) application. Being in a LEO orbit, HST has a 45-minute umbra period, during which spacecraft power requirements normally force the batteries into discharge, and a 60-minute sun period, which is available for battery recharge. The intent of this paper is to address the issue of NiH2 battery reliability and how battery capacity degradation can impact scheduling of a Servicing Mission to bring replacement batteries to HST, and extend mission life till deployment of Next Generation Space Telescope (NGST), planned for 2008 at the earliest.

Description: This final report presents the Graduate Student Research Program (GSRP) work Mr. Morrell was able to complete as a summer intern at NASA MSFS during the summer of 2001, and represents work completed from inception through project termination. The topics include: 1) NASA TD40 Organization; 2) Combustion Physics Lab; 3) Advanced Hydrocarbon Fuels; 4) GSRP Summer Tasks; 5) High Pressure Facility Installation; 6) High Pressure Combustion Issues; 7) High Energy Density Matter (HEDM) Hydrocarbons; and 8) GSRP Summer Intern Summary.

Description: A LOX/GH2 swirl injector was designed for a 10:1 propellant throttling range. To accomplish this, a dual LOX manifold was used feeding a single common vortex chamber of the swirl element. Hot-fire experiments were conducted for rocket chamber pressures from 80 to 800 psia at a mixture ratio of nominally 6.0 using steady flow, single-point-per-firing cases as well as dynamic throttling conditions. Low frequency (mean) and high frequency (fluctuating) pressure transducer data, flow meter measurements, and Raman spectroscopy images for mixing information were obtained. The injector design, experimental setup, low frequency pressure data, and injector performance analysis are presented. C* efficiency was very high (approx. 100%) at the middle of the throttleable range with somewhat lower performance at the high and low ends. From the analysis of discreet steady state operating conditions, injector pressure drop was slightly higher than predicted with an inviscid analysis, but otherwise agreed well across the design throttling range. Dynamic throttling of this injector was attempted with marginal success due to the immaturity of the throttling control system. Although the targeted mixture ratio of 6.0 was not maintained throughout the dynamic throttling profile, the injector behaved well over the wide range of conditions.

Description: The objective of the current effort is to provide a computational framework for design and analysis of the entire fuel supply system of a liquid rocket engine, including high-fidelity unsteady turbopump flow analysis. This capability is needed to support the design of pump sub-systems for advanced space transportation vehicles that are likely to involve liquid propulsion systems. To date, computational tools for design/analysis of turbopump flows are based on relatively lower fidelity methods. An unsteady, three-dimensional viscous flow analysis tool involving stationary and rotational components for the entire turbopump assembly has not been available for real-world engineering applications. The present effort provides developers with information such as transient flow phenomena at start up, and nonuniform inflows, and will eventually impact on system vibration and structures. In the proposed paper, the progress toward the capability of complete simulation of the turbo-pump for a liquid rocket engine is reported. The Space Shuttle Main Engine (SSME) turbo-pump is used as a test case for evaluation of the hybrid MPI/Open-MP and MLP versions of the INS3D code. CAD to solution auto-scripting capability is being developed for turbopump applications. The relative motion of the grid systems for the rotor-stator interaction was obtained using overset grid techniques. Unsteady computations for the SSME turbo-pump, which contains 114 zones with 34.5 million grid points, are carried out on Origin 3000 systems at NASA Ames Research Center. Results from these time-accurate simulations with moving boundary capability are presented along with the performance of parallel versions of the code.

Description: NASA Glenn Research Center is sponsoring efforts to develop technology for high-performance, high-density, low-freezing point, low-hazards monopropellant systems. The program is focused on a family of monopropellant formulations composed of an aqueous solution of hydroxylammonium nitrate (HAN) and a fuel component. HAN-based monopropellants offer significant mass and volume savings to small (less than 100 kg) satellite for orbit raising and on-orbit propulsion applications. The low-hazards characteristics of HAN-based monopropellants make them attractive for applications where ground processing costs are a significant concern. A 1-lbf thruster has been demonstrated to a 20-kg satellite orbit insertion duty cycle, using a formulation compatible with currently available catalysts. To achieve specific impulse levels above those of hydrazine, catalyst materials that can withstand the high-temperature, corrosive combustion environment of HAN-based monopropellants have to be developed. There also needs to be work done to characterize propellant properties, burning behavior, and material compatibility. NASA is coordinating their monopropellant efforts with those of the United States Air Force.

Description: This paper describes the concept and preliminary component testing of a gas-cooled, UN-fueled, pin-type reactor which uses He/Xe gas that goes directly into a recuperated Brayton system to produce electricity for nuclear electric propulsion. This Direct-Drive Gas-Cooled Reactor (DDG) is designed to be subcritical under water or wet- sand immersion in case of a launch accident. Because the gas-cooled reactor can directly drive the Brayton turbomachinery, it is possible to configure the system such that there are no external surfaces or pressure boundaries that are refractory metal, even though the gas delivered to the turbine is 1144 K. The He/Xe gas mixture is a good heat transport medium when flowing, and a good insulator when stagnant. Judicious use of stagnant cavities as insulating regions allows transport of the 1144-K gas while keeping all external surfaces below 900 K. At this temperature super-alloys (Hastelloy or Inconel) can be used instead of refractory metals. Super-alloys reduce the technology risk because they are easier to fabricate than refractory metals, we have a much more extensive knowledge base on their characteristics, and, because they have a greater resistance to oxidation, system testing is eased. The system is also relatively simple in its design: no additional coolant pumps, heat exchanger, or freeze-thaw systems are required. Key to success of this concept is a good knowledge of the heat transfer between the fuel pins and the gas, as well as the pressure drop through the system. This paper describes preliminary testing to obtain this key information, as well as experience in demonstrating electrically heated testing of simulated reactor components.

Description: One of the major advantages of using P-B11 fusion fuel is that the reaction produces only charged particles in the form of three alpha particles and no neutrons. A fusion concept that lends itself to this fuel cycle is the Magnetically Insulated Inertial Confinement Fusion (MICF) reactor whose distinct advantage lies in the very strong magnetic field that is created when an incident particle (or laser) beam strikes the inner wall of the target pellet. This field serves to thermally insulate the hot plasma from the metal wall thereby allowing thc plasma to burn for a long time and produce a large energy magnification. If used as a propulsion device, we propose using antiprotons to drive the system which we show to be capable of producing very large specific impulse and thrust. By way of validating the confinement propenies of MICF we will address a proposed experiment in which pellets coated with P-B11 fuel at the appropriate ratio will be zapped by a beam of antiprotons that enter the target through a hole. Calculations showing the density and temperature of the generated plasma along with the strength of the magnetic field and other properties of the system will be presented and discussed.

Description: An experimental Gasdynamic Mirror or GDM device has been constructed at the NASA Marshall Space Flight Center to provide an initial assessment of the applicability of this technology for propulsion systems. This paper presents the first experimental results obtained from the machine and an analysis of the types of plasma instabilities likely to be encountered. It is intended that this device operate at higher plasma densities and with much larger L/D ratios than previous mirror machines. The high L/D ratio minimizes to a large extent certain magnetic curvature effects which lead to plasma instabilities causing a loss of plasma confinement. The high plasma density results in the plasma behaving much more like a conventional fluid with a mean free path shorter than the length of the device. This characteristic helps reduce problems associated with "loss cone" microinstabilities. The device has been constructed to allow a considerable degree of flexibility in its configuration thus permitting the experiment to grow over time without necessitating a great deal of additional fabrication.

Description: This paper addresses the progress of technology development of a laser ignition system at NASA Marshall Space Flight Center (MSFC). The first two years of the project focus on comprehensive assessments and evaluations of a novel dual-pulse laser concept, flight- qualified laser system, and the technology required to integrate the laser ignition system to a rocket chamber. With collaborations of the Department of Energy/Los Alamos National Laboratory (LANL) and CFD Research Corporation (CFDRC), MSFC has conducted 26 hot fire ignition tests with lab-scale laser systems. These tests demonstrate the concept feasibility of dual-pulse laser ignition to initiate gaseous oxygen (GOX)/liquid kerosene (RP-1) combustion in a rocket chamber. Presently, a fiber optic- coupled miniaturized laser ignition prototype is being implemented at the rocket chamber test rig for future testing. Future work is guided by a technology road map that outlines the work required for maturing a laser ignition system. This road map defines activities for the next six years, with the goal of developing a flight-ready laser ignition system.

Description: Pulse detonation rocket engines (PDREs) offer potential performance improvements over conventional designs, but represent a challenging modellng task. A simplified model for an idealized, straight-tube, single-shot PDRE blowdown process and thrust determination is described and implemented. In order to form an assessment of the accuracy of the model, the flowfield time history is compared to experimental data from Stanford University. Parametric Studies of the effect of mixture stoichiometry, initial fill temperature, and blowdown pressure ratio on the performance of a PDRE are performed using the model. PDRE performance is also compared with a conventional steady-state rocket engine over a range of pressure ratios using similar gasdynamic assumptions.

Description: Preliminary results are presented for two methods to approximate the mission performance of high specific impulse high specific power vehicles. The first method is based on an analytical approximation derived by Williams and Shepherd and can be used to approximate mission performance to outer planets and interstellar space. The second method is based on a parametric analysis of trajectories created using the well known trajectory optimization code, VARITOP. This parametric analysis allows the reader to approximate payload ratios and optimal power requirements for both one-way and round-trip missions. While this second method only addresses missions to and from Jupiter, future work will encompass all of the outer planet destinations and some interstellar precursor missions.

Description: Previous efforts to develop power electronics for Hall thruster systems have targeted the 1 to 5 kW power range and an output voltage of approximately 300 V. New Hall thrusters are being developed for higher power, higher specific impulse, and multi-mode operation. These thrusters require up to 50 kW of power and a discharge voltage in excess of 600 V. Modular power supplies can process more power with higher efficiency at the expense of complexity. A 1 kW discharge power module was designed, built and integrated with a Hall thruster. The breadboard module has a power conversion efficiency in excess of 96 percent and weighs only 0.765 kg. This module will be used to develop a kW, multi-kW, and high voltage power processors.

Description: The 17th Space Photovoltaic Research and Technology (SPRAT XVII) Conference was held September 11-13, 2001, at the Ohio Aerospace Institute (OAI) in Cleveland, Ohio. The SPRAT conference, hosted by the Photovoltaic and Space Environments Branch of the NASA Glenn Research Center, brought together representatives of the space photovoltaic community from around the world to share the latest advances in space solar technology. This year's conference continued to build on many of the trends shown in SPRAT XVI; the use of new high-efficiency cells for commercial use and the development of novel array concepts such as Boeing's Solar Tile concept. In addition, new information was presented on space environmental interactions with solar arrays.

Description: Outer solar system missions will have propulsion system lifetime requirements well in excess of that which can be satisfied by ion thrusters utilizing conventional hollow cathode technology. To satisfy such mission requirements, other technologies must be investigated. One possible approach is to utilize electrodeless plasma production schemes. Such an approach has seen low power application less than 1 kW on earth-space spacecraft such as ARTEMIS which uses the rf thruster the RIT 10 and deep space missions such as MUSES-C which will use a microwave ion thruster. Microwave and rf thruster technologies are compared. A microwave-based ion thruster is investigated for potential high power ion thruster systems requiring very long lifetimes.

Description: This presentation provides an overview of IEC (Inertial Electrostatic Confinement) research and experiments at NASA's Marshall Space Flight Center. Topics covered include: apparatus involvement, iec schematics, iec plasma images, iec deuterium experiments, thomson scattering, detector options and experiment results.

Description: NASA's Glenn Research Center has been selected to lead development of NASA's Evolutionary Xenon Thruster (NEXT) system. The central feature of the NEXT system is an electric propulsion thruster (EPT) that inherits the knowledge gained through the NSTAR thruster that successfully propelled Deep Space 1 to asteroid Braille and comet Borrelly, while significantly increasing the thruster power level and making improvements in performance parameters associated with NSTAR. The EPT concept under development has a 40 cm beam diameter, twice the effective area of the Deep-Space 1 thruster, while maintaining a relatively-small volume. It incorporates mechanical features and operating conditions to maximize the design heritage established by the flight NSTAR 30 cm engine, while incorporating new technology where warranted to extend the power and throughput capability. The NASA Hall thruster program currently supports a number of tasks related to high power thruster development for a number of customers including the Energetics Program (formerly called the Space-based Program), the Space Solar Power Program, and the In-space Propulsion Program. In program year 2002, two tasks were central to the NASA Hall thruster program: 1.) the development of a laboratory Hall thruster capable of providing high thrust at high power; 2.) investigations into operation of Hall thrusters at high specific impulse. In addition to these two primary thruster development activities, there are a number of other on-going activities supported by the NASA Hall thruster program, These additional activities are related to issues such as thruster lifetime and spacecraft integration.

Description: Rocket engine oscillating data can reveal many physical phenomena ranging from unsteady flow and acoustics to rotordynamics and structural dynamics. Because of this, engine diagnostics based on oscillation data should employ both signal analysis and physical modeling. This paper describes an approach to rocket engine oscillation diagnostics, types of problems encountered, and example problems solved. Determination of design guidelines and environments (or loads) from oscillating phenomena is required during initial stages of rocket engine design, while the additional tasks of health monitoring, incipient failure detection, and anomaly diagnostics occur during engine development and operation. Oscillations in rocket engines are typically related to flow driven acoustics, flow excited structures, or rotational forces. Additional sources of oscillatory energy are combustion and cavitation. Included in the example problems is a sampling of signal analysis tools employed in diagnostics. The rocket engine hardware includes combustion devices, valves, turbopumps, and ducts. Simple models of an oscillating fluid system or structure can be constructed to estimate pertinent dynamic parameters governing the unsteady behavior of engine systems or components. In the example problems it is shown that simple physical modeling when combined with signal analysis can be successfully employed to diagnose complex rocket engine oscillatory phenomena.

Description: This presentation provides an overview of material covered during 'Space Shuttle Fuel Feedliner Cracking Investigation MSFC Fluids Workshop' held November 19-21, 2002. Topics covered include: cracks on fuel feed lines of Orbiter space shuttles, fluid driven cracking analysis, liner structural modes, structural motion in a fluid, fluid borne drivers, three dimensional computational fluid dynamics models, fluid borne drivers from pumps, amplification mechanisms, flow parameter mapping, and flight engine flow map.

Description: Space Solar Power (SSP), combined with Wireless Power Transmission (WPT), offers the far-term potential to solve major energy problems on Earth. In this paper two basic WPT options, using radio waves and light waves, are considered for both long-term and near-term SSP applications. In the long-term, we aspire to beam energy to Earth from geostationary Earth orbit (GEO), or even further distances in space. Accordingly, radio- and light- wave WPT options are compared through a wide range of criteria, each showing certain strengths. In the near-term, we plan to beam power over more moderate distances, but still stretch the limits of today's technology. For the near-term, a 100 kWe-class 'Power Plug' Satellite and a 10 kWe-class Lunar Polar Solar Power outpost are considered as the first steps in using these WPT options for SSP. By using SSP and WPT technology in near-term space science and exploration missions, we gain experience needed for sound decisions in designing and developing larger systems to send power from Space to Earth.

Description: SAIC is currently developing the Electric Propulsion Interactions Code 'EPIC', an interactive computer tool that allows the construction of a 3-D spacecraft model, and the assessment of interactions between its subsystems and the plume from an electric thruster. EPIC unites different computer tools to address the complexity associated with the interaction processes. This paper describes the overall architecture and capability of EPIC including the physics and algorithms that comprise its various components. Results from selected modeling efforts of different spacecraft-thruster systems are also presented.

Description: This simulation involved a two-dimensional axisymmetric model of a full motor initial grain of the Reusable Solid Rocket Motor (RSRM) of the Space Transportation System (STS). It was conducted with CFD (computational fluid dynamics) commercial code FLUENT. This analysis was performed to: a) maintain continuity with most related previous analyses, b) serve as a non-vectored baseline for any three-dimensional vectored nozzles, c) provide a relatively simple application and checkout for various CFD solution schemes, grid sensitivity studies, turbulence modeling and heat transfer, and d) calculate nozzle convective heat transfer coefficients. The accuracy of the present results and the selection of the numerical schemes and turbulence models were based on matching the rocket ballistic predictions of mass flow rate, head end pressure, vacuum thrust and specific impulse, and measured chamber pressure drop. Matching these ballistic predictions was found to be good. This study was limited to convective heat transfer and the results compared favorably with existing theory. On the other hand, qualitative comparison with backed-out data of the ratio of the convective heat transfer coefficient to the specific heat at constant pressure was made in a relative manner. This backed-out data was devised to match nozzle erosion that was a result of heat transfer (convective, radiative and conductive), chemical (transpirating), and mechanical (shear and particle impingement forces) effects combined.

Description: A series of full-scale buckling tests were performed on the space shuttle Reusable Solid Rocket Motor (RSRM) cylinders. The tests were performed to determine the buckling capability of the cylinders and to provide data for analytical comparison. A nonlinear ANSYS Finite Element Analysis (FEA) model was used to represent and evaluate the testing. Analytical results demonstrated excellent correlation to test results, predicting the failure load within 5%. The analytical value was on the conservative side, predicting a lower failure load than was applied to the test. The resulting study and analysis indicated the important parameters for FEA to accurately predict buckling failure. The resulting method was subsequently used to establish the pre-launch buckling capability of the space shuttle system.

Description: Flight qualification of ion thrusters typically requires testing on the order of 10,000 hours. Extensive knowledge of wear mechanisms and rates is necessary to establish design confidence prior to long duration tests. Consequently, real-time erosion rate measurements offer the potential both to reduce development costs and to enhance knowledge of the dependency of component wear on operating conditions. Several previous studies have used laser induced fluorescence (LIF) to measure real-time, in situ erosion rates of ion thruster accelerator grids. Those studies provided only relative measurements of the erosion rate. In the present investigation, a molybdenum tube was resistively heated such that the evaporation rate yielded densities within the tube on the order of those expected from accelerator grid erosion. A pulsed UV laser was used to pump the ground state molybdenum at 345.64nm, and the non-resonant fluorescence at 550-nm was collected using a bandpass filter and a photomultiplier tube or intensified CCD array. The sensitivity of the fluorescence was evaluated to determine the limitations of the calibration technique. The suitability of the diagnostic calibration technique was assessed for application to ion engine erosion rate measurements.

Description: Anions that can form dissociative salts with Li(+) have been prepared and covalently attached to high surface area fumed silica. When blended with polyethylene oxide (PEO), the functionalized fumed silica suppresses the crystallization of the PEO, provides dimensional stability, and serves as a single ion conductor. Since functionalized fumed silica is easily dispersed in common polar solvents, it can be incorporated in both the polymer electrolyte and the electrodes.

Description: This presentation reports on the NASA project to develop a prototype for RS-83 engine designed for use on reusable launch vehicles (RLV). Topics covered include: program objectives, overview schedule, organizational chart, integrated systems engineering processes, requirement analysis, catastrophic engine loss, maintainability analysis tools, and prototype design analysis.

Description: To obtain a better understanding the response of the structural adhesives used in the Space Shuttle's Reusable Solid Rocket Motor (RSRM) nozzle, an extensive effort has been conducted to characterize in detail the failure properties of these adhesives. This effort involved the development of a failure model that includes the effects of multi-axial loading, temperature, and time. An understanding of the effects of these parameters on the failure of the adhesive is crucial to the understanding and prediction of the safety of the RSRM nozzle. This paper documents the use of this newly developed multi-axial, temperature, and time (MATT) dependent failure model for modeling failure for the adhesives TIGA 321, EA913NA, and EA946. The development of the mathematical failure model using constant load rate normal and shear test data is presented. Verification of the accuracy of the failure model is shown through comparisons between predictions and measured creep and multi-axial failure data. The verification indicates that the failure model performs well for a wide range of conditions (loading, temperature, and time) for the three adhesives. The failure criterion is shown to be accurate through the glass transition for the adhesive EA946. Though this failure model has been developed and evaluated with adhesives, the concepts are applicable for other isotropic materials.

Description: The COBRA (CO-Optimized Booster for Reusable Applications) project include the following: 1. COBRA main engine project team. 2. COBRA and RLX cycles selected. 3. COBRA proto-type engine approach enables mission success. 4. COBRA provides quick, low cost demo of cycle and technologies. 5. COBRA cycle I risk reduction supports. 6. Achieving engine safety. 6. RLX cycle I risk reduction supports. 7. Flight qualification. 9. Life extension engine testing.

Description: Low- or no-flow electron emitters are required for low-power electric thrusters, spacecraft plasma contactors, and electrodynamic tether systems to reduce or eliminate the need for propellant/expellant. Expellant-less neutralizers can improve the viability of very low-power colloid thrusters, field emission electric propulsion devices, ion engines, Hall thrusters, and gridded vacuum arc thrusters. The NASA Glenn Research Center (GRC) is evaluating ferroelectric emission (FEE) cathodes as zero expellant flow rate cathode sources for the applications listed above. At GRC, low voltage (100s to approx. 1500 V) operation of FEE cathodes is examined. Initial experiments, with unipolar, bipolar, and RF burst applied voltage, have produced current pulses 250 to 1000 ns in duration with peak currents of up to 2 A at voltages at or below 1500 V. In particular, FEE cathodes driven by RF burst voltages from 1400 to 2000 V peak to peak, at burst frequencies from 70 to 400 kHz, emitted average current densities from 0.1 to 0.7 A/sq cm. Pulse repeatability as a function of input voltage has been initially established. Reliable emission has been achieved in air background at pressures as high as 10(exp -6) Torr.

Description: With the demonstration of the NSTAR propulsion system on the Deep Space One mission, the range of the Discovery class of NASA missions can now be expanded. NSTAR lacks, however, sufficient performance for many of the more challenging Office of Space Science (OSS) missions. Recent studies have shown that NASA's Evolutionary Xenon Thruster (NEXT) ion propulsion system is the best choice for many exciting potential OSS missions including outer planet exploration and inner solar system sample returns. The NEXT system provides the higher power, higher specific impulse, and higher throughput required by these science missions.