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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 847 records · Page 47
All sourcesNASA NTRSDOE OSTI
NASA NTRS1991

Space transfer vehicle concepts and requirements study. Volume 1: Executive summary

A description of the study in terms of background, objectives, and issues is provided. NASA is currently studying new initiatives of space exploration involving both piloted and unpiloted missions to destinations throughout the solar system. Many of these missions require substantial improvements in launch vehicle and upper stage capabilities. This study provides a focused examination of the Space Transfer Vehicles (STV) required to perform these missions using the emerging national launch vehicle definition, the Space Station Freedom (SSF) definition, and the latest mission scenario requirements. The study objectives are to define preferred STV concepts capable of accommodating future exploration missions in a cost-effective manner, determine the technology development (if any) required to perform these missions, and develop a decision database of various programmatic approaches for the development of the STV family of vehicles. Special emphasis was given to examining space basing (stationing reusable vehicles at a space station), examining the piloted lunar mission as a primary design mission, and restricting trade studies to the high-performance, near-term cryogenics (LO2/LH2) as vehicle propellant. The study progressed through three distinct 6-month phases. The first phase concentrated on supporting a NASA 3 month definition of exploration requirements (the '90-day study') and during this phase developed and optimized the space-based point-of-departure (POD) 2.5-stage lunar vehicle. The second phase developed a broad decision database of 95 different vehicle options and transportation architectures. The final phase chose the three most cost-effective architectures and developed point designs to carry to the end of the study. These reference vehicle designs are mutually exclusive and correspond to different national choices about launch vehicles and in-space reusability. There is, however, potential for evolution between concepts.

Weber, Gary A.↗
NASA NTRS1990

Propulsion needs for lunar/Mars missions

Viewgraphs on propulsion needs for lunar/Mars missions are presented. Topics covered include: lunar mission profile; lunar transfer vehicle/lunar excursion vehicle; launch vehicles for lunar missions; lunar outpost; Mars transfer operations; Mars mission vehicle in low earth orbit (LEO); mission vehicle commonality; mass comparison for reference missions; advanced propulsion; propulsion option size comparison; Mars transfer vehicle (MTV) propulsion option weights for mission favorable opportunities; propulsion options comparison; Mars transportation architecture options; tanker options for fully reusable systems; NTR 900 Isp staged tanks and engines; fully reusable cryogenic aerobraked system; NEP operated from high orbit; and nuclear safe orbit considerations.

Woodcock, Gordon R.↗
NASA NTRS1993

Thermal certification tests of Orbiter Thermal Protection System tiles coated with KSC coating slurries

Thermal tests of Orbiter thermal protection system (TPS) tiles, which were coated with borosilicate glass slurries fabricated at Kennedy Space Center (KSC), were performed in the Radiant Heat Test Facility and the Atmospheric Reentry Materials & Structures Evaluation Facility at Johnson Space Center to verify tile coating integrity after exposure to multiple entry simulation cycles in both radiant and convective heating environments. Eight high temperature reusable surface insulation (HRSI) tiles and six low temperature reusable surface insulation (LRSI) tiles were subjected to 25 cycles of radiant heat at peaked surface temperatures of 2300 F and 1200 F, respectively. For the LRSI tiles, an additional cycle at peaked surface temperature of 2100 F was performed. There was no coating crack on any of the HRSI specimens. However, there were eight small coating cracks (less than 2 inches long) on two of the six LRSI tiles on the 26th cycle. There was practically no change on the surface reflectivity, physical dimensions, or weight of any of the test specimens. There was no observable thermal-chemical degradation of the coating either. For the convective heat test, eight HRSI tiles were tested for five cycles at a surface temperature of 2300 F. There was no thermal-induced coating crack on any of the test specimens, almost no change on the surface reflectivity, and no observable thermal-chemical degradation with an exception of minor slumping of the coating under painted TPS identification numbers. The tests demonstrated that KSC's TPS slurries and coating processes meet the Orbiter's thermal specification requirements.

Milhoan, James D.↗
NASA NTRS1991

Examination of the effect of propulsion system performance variables on the life prediction for the SSME Lox Post

The main objective is to calculate damage evolution at the critical location of selected components of reusable space propulsion systems as the function of their operating environments and to relate the evolution of damage at the critical locations to the performance variables of the propulsion system using advanced constitutive and damage models. The models so developed will be oriented toward use in an advanced diagnostic/prognostic health monitoring system for reusable propulsion systems. The components to be modelled include a Lox Post (injector element), a thrust chamber, and a turbine blade. The study conducted on the Lox Post is reported.

Russell, D. A.↗
NASA NTRS1993

Damage-mitigating control of space propulsion systems for high performance and extended life

Calculations are presented showing that a substantial improvement in service life of a reusable rocket engine can be achieved by an insignificant reduction in the system dynamic performance. The paper introduces the concept of damage mitigation and formulates a continuous-time model of fatigue damage dynamics. For control of complex mechanical systems, damage prediction and damage mitigation are carried out based on the available sensory and operational information such that the plant can be inexpensively maintained and safely and efficiently steered under diverse operating conditions. The results of simulation experiments are presented for transient operations of a reusable rocket engine.

Ray, Asok↗
NASA NTRS1993

A ceramic matrix composite thermal protection system for hypersonic vehicles

The next generation of hypersonic vehicles (NASP, SSTO) that require reusable thermal protection systems will experience acreage surface temperatures in excess of 1100 C. More important, they will experience a more severe physical environment than the Space Shuttle due to non-pristine launching and landing conditions. As a result, maintenance, inspection, and replacement factors must be more thoroughly incorporated into the design of the TPS. To meet these requirements, an advanced thermal protection system was conceived, designated 'TOPHAT'. This system consists of a toughened outer ceramic matrix composite (CMC) attached to a rigid reusable surface insulator (RSI) which is directly bonded to the surface. The objective of this effort was to evaluate this concept in an aeroconvective environment, to determine the effect of impacts to the CMC material, and to compare the results with existing thermal protection systems.

Riccitiello, Salvatore R.↗
NASA NTRS1990

Automated Software Development Workstation (ASDW)

Software development is a serious bottleneck in the construction of complex automated systems. An increase of the reuse of software designs and components has been viewed as a way to relieve this bottleneck. One approach to achieving software reusability is through the development and use of software parts composition systems. A software parts composition system is a software development environment comprised of a parts description language for modeling parts and their interfaces, a catalog of existing parts, a composition editor that aids a user in the specification of a new application from existing parts, and a code generator that takes a specification and generates an implementation of a new application in a target language. The Automated Software Development Workstation (ASDW) is an expert system shell that provides the capabilities required to develop and manipulate these software parts composition systems. The ASDW is now in Beta testing at the Johnson Space Center. Future work centers on responding to user feedback for capability and usability enhancement, expanding the scope of the software lifecycle that is covered, and in providing solutions to handling very large libraries of reusable components.

Fridge, Ernie↗
NASA NTRS1993

Applying design principles to fusion reactor configurations for propulsion in space

The application of fusion power to space propulsion requires rethinking the engineering-design solution to controlled-fusion energy. Whereas the unit cost of electricity (COE) drives the engineering-design solution for utility-based fusion reactor configurations; initial mass to low earth orbit (IMLEO), specific jet power (kW(thrust)/kg(engine)), and reusability drive the engineering-design solution for successful application of fusion power to space propulsion. We applied three design principles (DP's) to adapt and optimize three candidate-terrestrial-fusion-reactor configurations for propulsion in space. The three design principles are: provide maximum direct access to space for waste radiation, operate components as passive radiators to minimize cooling-system mass, and optimize the plasma fuel, fuel mix, and temperature for best specific jet power. The three candidate terrestrial fusion reactor configurations are: the thermal barrier tandem mirror (TBTM), field reversed mirror (FRM), and levitated dipole field (LDF). The resulting three candidate space fusion propulsion systems have their IMLEO minimized and their specific jet power and reusability maximized. We performed a preliminary rating of these configurations and concluded that the leading engineering-design solution to space fusion propulsion is a modified TBTM that we call the Mirror Fusion Propulsion System (MFPS).

Carpenter, Scott A.↗
NASA NTRS1993

Fundamental rocket injector/spray programs at the Phillips Laboratory

The performance and stability of liquid rocket engines is determined to a large degree by atomization, mixing, and combustion processes. Control over these processes is exerted through the design of the injector. Injectors in liquid rocket engines are called upon to perform many functions. They must first of all mix the propellants to provide suitable performance in the shortest possible length. For main injectors, this is driven by the tradeoff between the combustion chamber performance, stability, efficiency, and its weight and cost. In gas generators and preburners, however, it is also driven by the possibility of damage to downstream components, for example piping and turbine blades. This can occur if unburned fuel and oxidant later react to create hot spots. Weight and cost considerations require that the injector design be simple and lightweight. For reusable engines, the injectors must also be durable and easily maintained. Suitable atomization and mixing must be produced with as small a pressure drop as possible, so that the size and weight of pressure vessels and turbomachinery can be minimized. However, the pressure drop must not be so small as to promote feed system coupled instabilities. Another important function of the injectors is to ensure that the injector face plate and the chamber and nozzle walls are not damaged. Typically this requires reducing the heat transfer to an acceptable level and also keeping unburned oxygen from chemically attacking the walls, particularly in reusable engines. Therefore the mixing distribution is often tailored to be fuel-rich near the walls. Wall heat transfer can become catastrophically damaging in the presence of acoustic instabilities, so the injector must prevent these from occurring at all costs. In addition to acoustic stability (but coupled with it), injectors must also be kinetically stable. That is, the flame itself must maintain ignition in the combustion chamber. This is not typically a problem with main injectors, but can be a consideration in preburners, where the desire to keep turbine inlet temperatures as cool as possible can make it advantageous for the preburners to operate as far from stoichiometry as can be tolerated. For some missions such as single stage to orbit, all of the above requirements must be maintained over a throttleable range, for example 5:1 to 10:1. Finally, the injectors must be ignitable during startup where pressures and temperatures are far from design conditions, and ignition transients must be minimized in order to avoid damage to engine components. In order to satisfy these various constraints, the injector designer must be able to perform design tradeoff studies, and it is important that this be done with minimal time and costs. In fact, it can easily be argued that reducing engine development time and costs is essential to maintaining U.S. competitiveness in space. The Propulsion Directorate of the Phillips Laboratory has invested in a number of programs to advance liquid rocket engine technology, and several of these are directed at improving design tools for liquid rocket injectors. The purpose of the presentation will be to describe some of these latter programs.

Talley, D. G.↗
NASA NTRS1993

Technology requirements for affordable single-stage rocket launch vehicles

A number of manned Earth-to-orbit (ETO) vehicle options for replacing or complementing the current Space Transportation System are being examined under the Advanced Manned Launch System (AMLS) study. The introduction of a reusable single-stage vehicle (SSV) into the U.S. launch vehicle fleet early in the next century could greatly reduce ETO launch costs. As a part of the AMLS study, the conceptual design of an SSV using a wide variety of enhancing technologies has recently been completed and is described in this paper. This paper also identifies the major enabling and enhancing technologies for a reusable rocket-powered SSV and provides examples of the mission payoff potential of a variety of important technologies. This paper also discusses the impact of technology advancements on vehicle margins, complexity, and risk, all of which influence the total system cost.

Stanley, Douglas O.↗
NASA NTRS1994

Test plan. GCPS task 7, subtask 7.1: IHM development

The overall objective of Task 7 is to identify cost-effective life cycle integrated health management (IHM) approaches for a reusable launch vehicle's primary structure. Acceptable IHM approaches must: eliminate and accommodate faults through robust designs, identify optimum inspection/maintenance periods, automate ground and on-board test and check-out, and accommodate and detect structural faults by providing wide and localized area sensor and test coverage as required. These requirements are elements of our targeted primary structure low cost operations approach using airline-like maintenance by exception philosophies. This development plan will follow an evolutionary path paving the way to the ultimate development of flight-quality production, operations, and vehicle systems. This effort will be focused on maturing the recommended sensor technologies required for localized and wide area health monitoring to a technology readiness level (TRL) of 6 and to establish flight ready system design requirements. The following is a brief list of IHM program objectives: design out faults by analyzing material properties, structural geometry, and load and environment variables and identify failure modes and damage tolerance requirements; design in system robustness while meeting performance objectives (weight limitations) of the reusable launch vehicle primary structure; establish structural integrity margins to preclude the need for test and checkout and predict optimum inspection/maintenance periods through life prediction analysis; identify optimum fault protection system concept definitions combining system robustness and integrity margins established above with cost effective health monitoring technologies; and use coupons, panels, and integrated full scale primary structure test articles to identify, evaluate, and characterize the preferred NDE/NDI/IHM sensor technologies that will be a part of the fault protection system.

Greenberg, H. S.↗
NASA NTRS1994

An operations and command systems for the extreme ultraviolet explorer

About 40% of the budget of a scientific spacecraft mission is usually consumed by Mission Operations & Data Analysis (MO&DA) with MO driving these costs. In the current practice, MO is separated from spacecraft design and comes in focus relatively late in the mission life cycle. As a result, spacecraft may be designed that are very difficult to operate. NASA centers have extensive MO expertise but often lessons learned in one mission are not exploited for other parallel or future missions. A significant reduction of MO costs is essential to ensure a continuing and growing access to space for the scientific community. We are addressing some of these issues with a highly automated payload operations and command system for an existing mission, the Extreme Ultraviolet Explorer (EUVE). EUVE is currently operated jointly by the Goddard Space Flight Center (GSFC), responsible for spacecraft operations, and the Center for Extreme Ultraviolet Astrophysics (CEA) of the University of California, Berkeley, which controls the telescopes and scientific instruments aboard the satellite. The new automated system is being developed by a team including personnel from the NASA Ames Research Center (ARC), the Jet Propulsion Laboratory (JPL) and the Center for EUV Astrophysics (CEA). An important goal of the project is to provide AI-based technology that can be easily operated by nonspecialists in AI. Another important goal is the reusability of the techniques for other missions. Models of the EUVE spacecraft need to be built both for planning/scheduling and for monitoring. In both cases, our modeling tools allow the assembly of a spacecraft model from separate sub-models of the various spacecraft subsystems. These sub-models are reusable; therefore, building mission operations systems for another small satellite mission will require choosing pre-existing modules, reparametrizing them with respect to the actual satellite telemetry information, and reassembling them in a new model. We briefly describe the EUVE mission and indicate why it is particularly suitable for the task. Then we briefly outline our current work in mission planning/scheduling and spacecraft and instrument health monitoring.

Muscettola, Nicola↗
NASA NTRS1995

Robust Rocket Engine Concept

The potential for a revolutionary step in the durability of reusable rocket engines is made possible by the combination of several emerging technologies. The recent creation and analytical demonstration of life extending (or damage mitigating) control technology enables rapid rocket engine transients with minimum fatigue and creep damage. This technology has been further enhanced by the formulation of very simple but conservative continuum damage models. These new ideas when combined with recent advances in multidisciplinary optimization provide the potential for a large (revolutionary) step in reusable rocket engine durability. This concept has been named the robust rocket engine concept (RREC) and is the basic contribution of this paper. The concept also includes consideration of design innovations to minimize critical point damage.

Lorenzo, Carl F.↗
NASA NTRS1995

Human Exploration and Settlement of the Moon Using LUNOX-Augmented NTR Propulsion

An innovative trimodal nuclear thermal rocket (NTR) concept is described which combines conventional liquid hydrogen (LH2)-cooled NTR, Brayton cycle power generation and supersonic combustion ramjet (scramjet) technologies. Known as the liquid oxygen (LOX) augmented NTR (LANTR), this concept utilizes the large divergent section of the NTR nozzle as an 'afterburner' into which LOX is injected and supersonically combusted with nuclear preheated hydrogen emerging from the LANTR's choked sonic throat--'scramjet propulsion in reverse.' By varying the oxygen-to-hydrogen mixture ratio (MR), the LANTR can operate over a wide range of thrust and specific impulse (Isp) values while the reactor core power level remains relatively constant. As the MR varies from zero to seven, the thrust-to-weight ratio for a 15 thousand pound force (klbf) NTR increases by approximately 440%--from 3 to 13--while the Isp decreases by only approximately 45%--from 940 to 515 seconds. This thrust augmentation feature of the LANTR means that 'big engine' performance can be obtained using smaller more affordable, easier to test NTR engines. 'Reoxidizing' the bipropellant LANTR system in low lunar orbit (LLO) with high density 'lunar-derived' LOX (LUNOX) enables a reusable, reduced size and mass lunar transfer vehicle (LTV) which can be deployed and resupplied using two 66 t-class Shuttle-derived launch vehicles. The reusable LANTR can also transport 200 to 300% more payload on each piloted round trip mission than an expendable 'all LH2' NTR system. As initial outposts grow to eventual lunar settlements and LUNOX production capacity increases, the LANTR concept can also enable a rapid 'commuter' shuttle capable of 36 to 24 hour 'one way' trips to the Moon and back with reasonable size vehicles and initial mass in low Earth orbit (IMLEO) requirements.

Borowski, Stanley K.↗
NASA NTRS1989

SSME component assembly and life management expert system

The space shuttle utilizes several rocket engine systems, all of which must function with a high degree of reliability for successful mission completion. The space shuttle main engine (SSME) is by far the most complex of the rocket engine systems and is designed to be reusable. The reusability of spacecraft systems introduces many problems related to testing, reliability, and logistics. Components must be assembled from parts inventories in a manner which will most effectively utilize the available parts. Assembly must be scheduled to efficiently utilize available assembly benches while still maintaining flight schedules. Assembled components must be assigned to as many contiguous flights as possible, to minimize component changes. Each component must undergo a rigorous testing program prior to flight. In addition, testing and assembly of flight engines and components must be done in conjunction with the assembly and testing of developmental engines and components. The development, testing, manufacture, and flight assignments of the engine fleet involves the satisfaction of many logistical and operational requirements, subject to many constraints. The purpose of the SSME Component Assembly and Life Management Expert System (CALMES) is to assist the engine assembly and scheduling process, and to insure that these activities utilize available resources as efficiently as possible.

Ali, M.↗
NASA NTRS1997

Evaluation of Thermal Control Coatings for Flexible Ceramic Thermal Protection Systems

This report summarizes the evaluation and testing of high emissivity protective coatings applied to flexible insulations for the Reusable Launch Vehicle technology program. Ceramic coatings were evaluated for their thermal properties, durability, and potential for reuse. One of the major goals was to determine the mechanism by which these coated blanket surfaces become brittle and try to modify the coatings to reduce or eliminate embrittlement. Coatings were prepared from colloidal silica with a small percentage of either SiC or SiB6 as the emissivity agent. These coatings are referred to as gray C-9 and protective ceramic coating (PCC), respectively. The colloidal solutions were either brushed or sprayed onto advanced flexible reusable surface insulation blankets. The blankets were instrumented with thermocouples and exposed to reentry heating conditions in the Ames Aeroheating Arc Jet Facility. Post-test samples were then characterized through impact testing, emissivity measurements, chemical analysis, and observation of changes in surface morphology. The results show that both coatings performed well in arc jet tests with backface temperatures slightly lower for the PCC coating than with gray C-9. Impact testing showed that the least extensive surface destruction was experienced on blankets with lower areal density coatings.

Kourtides, Demetrius↗
NASA NTRS1996

Affordable In-Space Transportation Phase 2: An Advanced Concepts Project

The Affordable In-Space Transportation (AIST) program was established by the NASA Office of Space Access to improve transportation and lower the costs from Low Earth Orbit (LEO) to Geostationary Earth Orbit (GEO) and beyond (to Lunar orbit, Mars orbit, inner solar system missions, and return to LEO). A goal was established to identify and develop radically innovative concepts for new upper stages for Reusable Launch Vehicles (RLV) and Highly Reusable Space Transportation (HRST) systems. New architectures and technologies are being identified which have the potential to meet a cost goal of $1,000 to $2,000 per pound for transportation to GEO and beyond for overall mission cost (including the cost to LEO). A Technical Interchange Meeting (TTM) was held on October 16 and 17, 1996 in Huntsville, Alabama to review previous studies, present advanced concepts and review technologies that could be used to meet the stated goals. The TIN4 was managed by NASA-Marshall Space Flight Center (MSFC) Advanced Concepts Office with Mr. Alan Adams providing TIM coordination. Mr. John C. Mankins of NASA Headquarters provided overall sponsorship. The University of Alabama in Huntsville (UAH) Propulsion Research Center hosted the TIM at the UAH Research Center. Dr. Clark Hawk, Center Director, was the principal investigator. Technical support was provided by Christensen Associates. Approximately 70 attendees were present at the meeting. This Executive Summary provides a record of the key discussions and results of the TIN4 in a summary for-mat. It incorporates the response to the following basic issues of the TDVL which addressed the following questions: 1. What are the cost drivers and how can they be reduced? 2. What are the operational issues and their impact on cost? 3. What is the current technology readiness level (TRL) and what will it take to reach TRL 6? 4. What are the key enabling technologies and sequence for their accomplishment? 5 . What is the proposed implementation time frame? See Appendix A for the TIM Agenda and Appendix C for the AIST Program Terms of Reference.

Source record↗
NASA NTRS1996

Affordable In-Space Transportation: An Advanced Concepts Project - Phase 2

The Affordable In-Space Transportation (AIST) program was established by the NASA Office of Space Access to improve transportation and lower the costs from Low Earth Orbit (LEO) to Geostationary Earth Orbit (GEO) and beyond (to Lunar orbit, Mars orbit, inner solar system missions, and return to LEO). A goal was established to identify and develop radically innovative concepts for new upper stages for Reusable Launch Vehicles (RLV) and Highly Reusable Space Transportation (HRST) systems. New architectures and technologies are being identified which have the potential to meet a cost goal of $1,000 to $2,000 per pound for transportation to GEO and beyond for overall mission cost (including the cost to LEO). A Technical Interchange Meeting (ITM) was held on October 16 and 17, 1996 in Huntsville, Alabama to review previous studies, present advanced concepts and review technologies that could be used to meet the stated goals. The TIM was managed by NASA-Mar-shaU Space Flight Center (MSFC) Advanced Concepts Office with Mr. Alan Adams providing TIM coordination. Mr. John C. Manidns of NASA Headquarters provided overall sponsorship. The University of Alabama in Huntsville (UAH) Propulsion Research Center hosted the TM at the UAH Research Center. Dr. Clark Hawk, Center Director, was the principal investigator. Technical support was provided by Christensen Associates. Approximately 70 attendees were present at the meeting. This Executive Summary provides a record of the key discussions and results of the TIM in a summary format. It incorporates the response to the following basic issues of the TPA, which addressed the following questions: 1. What are the cost drivers and how can they be reduced? 2. What are the operational issues and their impact on cost? What is the current Technology Readiness Level (TRL) and what will it take to reach TRL 6? 4. What are the key enabling technologies and sequence for their accomplishment? 5. What is the proposed implementation time frame? See Appendix A for the TIM Agenda and Appendix C for the AIST Program Terms of Reference.

Source record↗