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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 217 records · Page 12

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NASA NTRS1971

Absolute stability analysis of attitude control systems for large boosters.

A method for performing an absolute stability analysis of attitude control systems for large launch vehicles is presented. Absolute stability of these systems is shown to be of a limited extent. The regions of absolute stability are computed by using a quadratic Liapunov function. The function is chosen to provide additional information about the exponential property of absolute stability. A system model is used to illustrate the method.

Siljak, D. D.↗

NASA NTRS2016

Design and Flight Performance of the Orion Pre-Launch Navigation System

Launched in December 2014 atop a Delta IV Heavy from the Kennedy Space Center, the Orion vehicle's Exploration Flight Test-1 (EFT-1) successfully completed the objective to test the prelaunch and entry components of the system. Orion's pre-launch absolute navigation design is presented, together with its EFT-1 performance.

Zanetti, Renato↗

NASA NTRS2001

An Approach to Establishing System Benefits for Technologies In NASA's Spaceliner Investment Area

NASA's has established long term goals for access-to-space. The third generation launch systems are to be fully reusable and operational around 2025. The goals for the third generation launch system are to significantly reduce cost and improve safety over current systems. The Advanced Space Transportation Program Office (ASTP) at the NASA's Marshall Space Flight Center in Huntsville, AL has the agency lead to develop space transportation technologies. Within ASTP, under the Spaceliner Investment Area, third generation technologies are being pursued. The Spaceliner Investment Area's primary objective is to mature vehicle technologies to enable substantial increases in the design and operating margins of third generation RLVs (current Space Shuttle is considered the first generation RLV) by incorporating advanced propulsion systems, materials, structures, thermal protection systems, power, and avionics technologies. Advancements in design tools and better characterization of the operational environment will result in reduced design and operational variabilities leading to improvements in margins. Improvements in operational efficiencies will be obtained through the introduction of integrated vehicle health management, operations and range technologies. Investments in these technologies will enable the reduction in the high operational costs associated with today's vehicles by allowing components to operate well below their design points resulting in improved component operating life, reliability, and safety which in turn reduces both maintenance and refurbishment costs. The introduction of advanced technologies may enable horizontal takeoff by significantly reducing the takeoff weight and allowing use of existing infrastructure. This would be a major step toward the goal of airline-like operation. These factors in conjunction with increased flight rates, resulting from reductions in transportation costs, will result in significant improvements of future vehicles. The real-world problem is that resources are limited and technologies need to be prioritized to assure the resources are spent on technologies that provide the highest system level benefits. Toward that end, a systems approach is being taken to determine the benefits of technologies for the Spaceliner Investment Area. Technologies identified to be enabling will be funded. However, the other technologies will be funded based on their system's benefits. Since the final launch system concept will not be decided for many years, several vehicle concepts are being evaluated to determine technology benefits. Not only performance, but also cost and operability are being assessed. This will become an annual process to assess these technologies against their goals and the benefits to various launch systems concepts. The paper describes the system process, tools and concepts used to determine the technology benefits. Preliminary results will be presented along with the current technology investments that are being made by ASTP's Spaceliner Investment Area.

Hueter, Uwe↗

NASA NTRS2012

Developmental Testing of Electric Thrust Vector Control Systems for Manned Launch Vehicle Applications

This paper describes recent developmental testing to verify the integration of a developmental electromechanical actuator (EMA) with high rate lithium ion batteries and a cross platform extensible controller. Testing was performed at the Thrust Vector Control Research, Development and Qualification Laboratory at the NASA George C. Marshall Space Flight Center. Electric Thrust Vector Control (ETVC) systems like the EMA may significantly reduce recurring launch costs and complexity compared to heritage systems. Electric actuator mechanisms and control requirements across dissimilar platforms are also discussed with a focus on the similarities leveraged and differences overcome by the cross platform extensible common controller architecture.

Bates, Lisa B.↗

NASA NTRS1992

Integrated health monitoring and controls for rocket engines

Current research in intelligent control systems at the Lewis Research Center is described in the context of a functional framework. The framework is applicable to a variety of reusable space propulsion systems for existing and future launch vehicles. It provides a 'road map' technology development to enable enhanced engine performance with increased reliability, durability, and maintainability. The framework hierarchy consists of a mission coordination level, a propulsion system coordination level, and an engine control level. Each level is described in the context of the Space Shuttle Main Engine. The concept of integrating diagnostics with control is discussed within the context of the functional framework. A distributed real time simulation testbed is used to realize and evaluate the functionalities in closed loop.

Merrill, W. C.↗

NASA NTRS1992

Integrated health monitoring and controls for rocket engines

Merrill, W. C.↗

NASA NTRS1995

Optimal technology investment strategies for a reusable launch vehicle

Within the present budgetary environment, developing the technology that leads to an operationally efficient space transportation system with the required performance is a challenge. The present research focuses on a methodology to determine high payoff technology investment strategies. Research has been conducted at Langley Research Center in which design codes for the conceptual analysis of space transportation systems have been integrated in a multidisciplinary design optimization approach. The current study integrates trajectory, propulsion, weights and sizing, and cost disciplines where the effect of technology maturation on the development cost of a single stage to orbit reusable launch vehicle is examined. Results show that the technology investment prior to full-scale development has a significant economic payoff. The design optimization process is used to determine strategic allocations of limited technology funding to maximize the economic payoff.

Moore, A. A.↗

NASA NTRS2006

Development of the J-2X Engine for the Ares I Crew Launch Vehicle and the Ares V Cargo Launch Vehicle: Building on the Apollo Program for Lunar Return Missions

The United States (U.S.) Vision for Space Exploration directs NASA to develop two new launch vehicles for sending humans to the Moon, Mars, and beyond. In January 2006, NASA streamlined its hardware development approach for replacing the Space Shuttle after it is retired in 2010. Benefits of this approach include reduced programmatic and technical risks and the potential to return to the Moon by 2020, by developing the Ares I Crew Launch Vehicle (CLV) propulsion elements now, with full extensibility to future Ares V Cargo Launch Vehicle (CaLV) lunar systems. This decision was reached after the Exploration Launch Projects Office performed a variety of risk analyses, commonality assessments, and trade studies. The Constellation Program selected the Pratt & Whitney Rocketdyne J-2X engine to power the Ares I Upper Stage Element and the Ares V Earth Departure Stage. This paper narrates the evolution of that decision; describes the performance capabilities expected of the J-2X design, including potential commonality challenges and opportunities between the Ares I and Ares V launch vehicles; and provides a current status of J-2X design, development, and hardware testing activities. This paper also explains how the J-2X engine effort mitigates risk by building on the Apollo Program and other lessons lived to deliver a human-rated engine that is on an aggressive development schedule, with its first demonstration flight in 2012.

Snoddy, Jim↗

NASA NTRS2011

Maturation of Structural Health Management Systems for Solid Rocket Motors

Concepts of an autonomous and automated space-compliant diagnostic system were developed for conditioned-based maintenance (CBM) of rocket motors for space exploration vehicles. The diagnostic system will provide real-time information on the integrity of critical structures on launch vehicles, improve their performance, and greatly increase crew safety while decreasing inspection costs. Using the SMART Layer technology as a basis, detailed procedures and calibration techniques for implementation of the diagnostic system were developed. The diagnostic system is a distributed system, which consists of a sensor network, local data loggers, and a host central processor. The system detects external impact to the structure. The major functions of the system include an estimate of impact location, estimate of impact force at impacted location, and estimate of the structure damage at impacted location. This system consists of a large-area sensor network, dedicated multiple local data loggers with signal processing and data analysis software to allow for real-time, in situ monitoring, and longterm tracking of structural integrity of solid rocket motors. Specifically, the system could provide easy installation of large sensor networks, onboard operation under harsh environments and loading, inspection of inaccessible areas without disassembly, detection of impact events and impact damage in real-time, and monitoring of a large area with local data processing to reduce wiring.

Quing, Xinlin↗

NASA NTRS1999

[X-33 Launch and Landing Facilities]

Sverdrup is responsible for the design, construction and activation of the X-33 Flight Operations Center at Edwards Air Force Base and for providing assistance in activating the X-33 Landing Sites. The past year has seen the completion of the construction of the X-33 Flight Operations Center. Construction was completed in December of 1998, with systems checkout and testing continuing into early 1999. Integration of the site with LMCMS and other partner-supplied systems began in December and will continue through rollout of the X-33 vehicle. The construction of the X-33 Launch Complex has been performed within the Edwards AFB and Air Force Research Laboratory (AFRL) systems with no substantial interference to either parties. A high level of cooperation exists between Sverdrup, Edwards AFB, and the Air Force Research Laboratory in the areas of access, training, security, and operations. There have been no conflicts between programs that have not been accommodated. Development of the landing sites is progressing with many of the modifications necessary underway. GSE commitments are in place. The personnel training program developed by Sverdrup for persons entering the launch site construction areas, was modified by Lockheed for use in training and access control to the Center during flight operations to maximize safety and minimize intrusion upon the environment. Close cooperation between Sverdrup, the construction workers, and the environmental biologist permitted construction to proceed in a timely fashion without harm to the wildlife, in particular, the Desert Tortoise. Although the entire X-33 site encompasses approximately 50 acres including a new access road, only the areas directly impacted by the construction were cleared to minimize the impact on the environment. A total of about 30 acres was actually disturbed.

Source record↗

NASA NTRSDate not supplied

Artemis I Space Launch System Base Heat Shield Thermal Protection System Performance

The Space Launch System (SLS) Core Stage base heat shield experienced the highest external heating environments on the entire launch vehicle during Artemis I ascent flight. This result was consistent with design predictions. The base heat shield experiences P50 cork combustion dynamics at low altitudes, plume-induced recirculation at moderate altitudes and then in-space base flow physics out to Main Engine Cut-Off (MECO). The base heat shield thermal protection system (TPS) is composed of a P50 cork ablator which is bonded to a substrate. The heat shield protects the gimbal actuation system, RS-25 turbomachinery systems and other aft section sensitive components during ascent. This paper estimates the base heat shield TPS performance from Artemis I using flight data from the NASA Langley Research Center’s Scientifically Calibrated In-Flight Imagery (SCIFLI) Airborne Multispectral Imager (SAMI), development flight instrumentation (DFI) and other TPS recession flight measurements. Predictions from computational and ground test-derived engineering ablation models and observations are also applied. Since no base heat shield substrate thermocouple data were obtained for Artemis I, an estimate of the TPS performance data is derived here. This data assesses thermal margin of the SLS Core Stage base heat shield and best informs the Artemis II Crewed mission to the moon.

aerothermodynamics↗

NASA NTRS1966

Gemini launch vehicle development

This paper presents a brief description of the basic modifications made to the Titan II to adapt it to a Gemini launch vehicle (GLV), the ground rules under which they were made, how the principal systems were initially baselined, how they evolved, and how they have performed to date.

TITAN II ICBM↗

NASA NTRS1974

Apollo Saturn 511 effluent measurements from the Apollo 16 launch operations: An experiment

An experiment was performed in conjunction with the Apollo 16 launch to define operational and instrumentational problems associated with launch-vehicle exhaust effluent monitoring. Ground and airborne sampling were performed for CO, CO2, hydrocarbons, and particulates. Sampling systems included filter pads and photometers for particulates and whole-air grab samples for gases. Launch debris was identified in the particulate samples at ground level(taken immediately after launch) and in the airborne measurements (taken 40 to 50 minutes after launch approximately 40 km downwind of the pad). Operational problems were identified and included the need for higher instrumentation mobility and the need for real-time sampling instrumentation as opposed to collection-type samples such as the whole-air grab sample.

Gregory, G. L.↗

NASA NTRS1974

Orbit/launch vehicle tradeoff studies. Earth Observatory Satellite system definition study (EOS)

An evaluation of the Earth Observatory Satellite (EOS) design, performance, and cost factors which affect the choices of an orbit and a launch vehicle is presented. Primary emphasis is given to low altitude (300 to 900 nautical miles) land resources management applications for which payload design factors are defined. The subjects considered are: (1) a mission model, (2) orbit analysis and characterization, (3) characteristics and capabilities of candidate conventional launch vehicles, and space shuttle support. Recommendations are submitted for the EOS-A mission, the Single Multispectral Scanner payload, the Single Multispectral Scanner plus Thematic Mapper payload, the Dual Multispectral Scanner payload, and the Dual Multispectral Scanner plus Thematic Mapper payload.

Source record↗

NASA NTRS2004

Automating Structural Analysis of Spacecraft Vehicles

A major effort within NASA's vehicle analysis discipline has been to automate structural analysis and sizing optimization during conceptual design studies of advanced spacecraft. Traditional spacecraft structural sizing has involved detailed finite element analysis (FEA) requiring large degree-of-freedom (DOF) finite element models (FEM). Creation and analysis of these models can be time consuming and limit model size during conceptual designs. The goal is to find an optimal design that meets the mission requirements but produces the lightest structure. A structural sizing tool called HyperSizer has been successfully used in the conceptual design phase of a reusable launch vehicle and planetary exploration spacecraft. The program couples with FEA to enable system level performance assessments and weight predictions including design optimization of material selections and sizing of spacecraft members. The software's analysis capabilities are based on established aerospace structural methods for strength, stability and stiffness that produce adequately sized members and reliable structural weight estimates. The software also helps to identify potential structural deficiencies early in the conceptual design so changes can be made without wasted time. HyperSizer's automated analysis and sizing optimization increases productivity and brings standardization to a systems study. These benefits will be illustrated in examining two different types of conceptual spacecraft designed using the software. A hypersonic air breathing, single stage to orbit (SSTO), reusable launch vehicle (RLV) will be highlighted as well as an aeroshell for a planetary exploration vehicle used for aerocapture at Mars. By showing the two different types of vehicles, the software's flexibility will be demonstrated with an emphasis on reducing aeroshell structural weight. Member sizes, concepts and material selections will be discussed as well as analysis methods used in optimizing the structure. Analysis based on the HyperSizer structural sizing software will be discussed. Design trades required to optimize structural weight will be presented.

Hrinda, Glenn A.↗

NASA NTRS1972

Basic parameters for low thrust mission and system analysis.

Development of a basic set of normalized parameters to represent low-thrust system characteristics. The conditions which must be met to obtain trajectory and performance data which are invariant with the launch vehicle characteristics, the selected power, the specific mass of the low-thrust vehicle, and the efficiency of the ion propulsion system are derived. This normalized representation of mission characteristics frees the analyst from having to rerun the trajectories if any of the aforementioned parameters are varied. Examples pertaining to a low-thrust Jupiter flyby and comet Encke rendezvous are given to demonstrate the utility of the method of presentation.

Barber, T. A.↗

NASA NTRS2002

GTX Reference Vehicle Structural Verification Methods and Weight Summary

The design of a single-stage-to-orbit air breathing propulsion system requires the simultaneous development of a reference launch vehicle in order to achieve the optimal mission performance. Accordingly, for the GTX study a 300-lb payload reference vehicle was preliminary sized to a gross liftoff weight (GLOW) of 238,000 lb. A finite element model of the integrated vehicle/propulsion system was subjected to the trajectory environment and subsequently optimized for structural efficiency. This study involved the development of aerodynamic loads mapped to finite element models of the integrated system in order to assess vehicle margins of safety. Commercially available analysis codes were used in the process along with some internally developed spread-sheets and FORTRAN codes specific to the GTX geometry for mapping of thermal and pressure loads. A mass fraction of 0.20 for the integrated system dry weight has been the driver for a vehicle design consisting of state-of-the-art composite materials in order to meet the rigid weight requirements. This paper summarizes the methodology used for preliminary analyses and presents the current status of the weight optimization for the structural components of the integrated system.

Hunter, J. E.↗

NASA NTRS2002

GTX Reference Vehicle Structural Verification Methods and Weight Summary

The design of a single-stage-to-orbit air breathing propulsion system requires the simultaneous development of a reference launch vehicle in order to achieve the optimal mission performance. Accordingly, for the GTX study a 300-lb payload reference vehicle was preliminarily sized to a gross liftoff weight (GLOW) of 238,000 lb. A finite element model of the integrated vehicle/propulsion system was subjected to the trajectory environment and subsequently optimized for structural efficiency. This study involved the development of aerodynamic loads mapped to finite element models of the integrated system in order to assess vehicle margins of safety. Commercially available analysis codes were used in the process along with some internally developed spreadsheets and FORTRAN codes specific to the GTX geometry for mapping of thermal and pressure loads. A mass fraction of 0.20 for the integrated system dry weight has been the driver for a vehicle design consisting of state-of-the-art composite materials in order to meet the rigid weight requirements. This paper summarizes the methodology used for preliminary analyses and presents the current status of the weight optimization for the structural components of the integrated system.

Hunter, J. E.↗