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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 289 records · Page 16
All sourcesNASA NTRSDOE OSTI
NASA NTRS2005

A Reliability Model to Assess Reliability of Shuttle Derived Launch Vehicle and Next Generation Vehicles

One important thrust in NASA s space exploration vision is to improve the safety and reliability of future launch vehicles. A quantitative reliability model is an essential technical tool to evaluate launch vehicle reliability and to enhance launch vehicle configuration down selection and detailed designs. A quantitative reliability model that supports NASA's thrust has been developed and implemented in an MS Excel spreadsheet. The model addresses key reliability parameters, and provides for sensitivity analysis of various vehicle designs. This paper presents the specific elements of the model and examples of selected sensitivity analysis results. Reliability input parameters in the model include the following: 1. Main propulsion element failure probability; 2. Propulsion element catastrophic failure fraction (Cf). This is defined as the ratio of probability of uncontained failures over sum of probability of uncontained failures and contained failures); 3. Engine-out design vs. no-engine-out design; 4. Engine with redlines vs. no redlines; 5. Different levels of health management system implementation; 6. Launch Escape System (LES) reliability; and 7. Power level correlation with reliability. The sensitivity results will illustrate the impact of each of the above parameters on several reliability metrics: Loss of Mission (LOM), LOV (Loss of Vehicle) and Loss of Crew (LOC).

Huang, Zhao↗
NASA NTRS1993

A shadowgraph study of two proposed Shuttle-C launch vehicle configurations

A shadowgraph study concerning two of the proposed Shuttle-C launch vehicle configurations are presented. These shadowgraphs were obtained from a wind tunnel test performed in Marshall Space Flight Center's 14-in trisonic wind tunnel at various angles-of-attack and roll angles over the Mach range of 0.6 to 4.96. Variations in payload bay length were also evaluated. Major flow field phenomena can easily be seen in the shadowgraphs. Shadowgraphs are a valuable resource. They are used in the analysis of the external flow conditions the launch vehicle encounters through the ascent stage of flight. Subsequent reports will contain shadowgraph studies for other launch vehicle configurations also tested in the Marshall Space Flight Center's 14-in trisonic wind tunnel.

Springer, A. M.↗
NASA NTRS2017

Analysis of Rawinsonde Spatial Separation for Space Launch Vehicle Applications at the Eastern Range

Space launch vehicles develop day-of-launch steering commands based upon the upper-level atmospheric environments in order to alleviate wind induced structural loading and optimize ascent trajectory. Historically, upper-level wind measurements to support launch operations at the National Aeronautics and Space Administration's (NASA's) Kennedy Space Center co-located on the United States Air Force's Eastern Range (ER) at the Cape Canaveral Air Force Station use high-resolution rawinsondes. One inherent limitation with rawinsondes consists of taking approximately one hour to generate a vertically complete wind profile. Additionally, rawinsonde drift during ascent by the ambient wind environment can result in the balloon being hundreds of kilometers down range, which results in questioning whether the measured winds represent the wind environment the vehicle will experience during ascent. This paper will describe the use of balloon profile databases to statistically assess the drift distance away from the ER launch complexes during rawinsonde ascent as a function of season and discuss an alternative method to measure upper level wind environments in closer proximity to the vehicle trajectory launching from the ER.

Decker, Ryan K.↗
NASA NTRS2010

A Methodology for Mapping Launch Vehicle Buffet Loads

Buffet loads represent the primary source of high frequency loading for launch vehicles during the ascent portion of flight. Currently, experimental techniques establish the nature of buffeting using a rigid scale model of the vehicle. The buffet forcing functions resulting from such tests are then applied to reduced finite-element models of the full-scale vehicle to determine the response and consequent loading. This paper discusses the techniques required to translate model-derived, empirical buffet forcing functions into responses for the full-scale launch vehicle, as used to determine the buffet loading for NASA's Ares I launch vehicle.

Schwarz, Jordan B.↗
NASA NTRS2011

Mutual Coupling of Internal Transmit/Receive Pair in Launch Vehicle Fairing Model Using WIPL-D

Evaluating the fairing Radio Frequency (RF) Environment within the launch vehicle payload fairing cavity due to internal transmitters is an issue for the spacecraft and launch vehicle industry. This paper provides an effective approach for launch vehicle fairing evaluation of power reception and field distribution due to internal transmitters. A commercial electromagnetic computational tool, WIPL-D is applied in this study for test data comparison.

Trout, Dawn H.↗
NASA NTRS2014

Airborne Simulation of Launch Vehicle Dynamics

In this paper we present a technique for approximating the short-period dynamics of an exploration-class launch vehicle during flight test with a high-performance surrogate aircraft in relatively benign endoatmospheric flight conditions. The surrogate vehicle relies upon a nonlinear dynamic inversion scheme with proportional-integral feedback to drive a subset of the aircraft states into coincidence with the states of a time-varying reference model that simulates the unstable rigid body dynamics, servodynamics, and parasitic elastic and sloshing dynamics of the launch vehicle. The surrogate aircraft flies a constant pitch rate trajectory to approximate the boost phase gravity-turn ascent, and the aircraft's closed-loop bandwidth is sufficient to simulate the launch vehicle's fundamental lateral bending and sloshing modes by exciting the rigid body dynamics of the aircraft. A novel control allocation scheme is employed to utilize the aircraft's relatively fast control effectors in inducing various failure modes for the purposes of evaluating control system performance. Sufficient dynamic similarity is achieved such that the control system under evaluation is optimized for the full-scale vehicle with no changes to its parameters, and pilot-control system interaction studies can be performed to characterize the effects of guidance takeover during boost. High-fidelity simulation and flight test results are presented that demonstrate the efficacy of the design in simulating the Space Launch System (SLS) launch vehicle dynamics using NASA Dryden Flight Research Center's Full-scale Advanced Systems Testbed (FAST), a modified F/A-18 airplane, over a range of scenarios designed to stress the SLS's adaptive augmenting control (AAC) algorithm.

Gilligan, Eric T.↗
NASA NTRS1995

Launch vehicle flight control augmentation using smart materials and advanced composites (CDDF Project 93-05)

The Marshall Space Flight Center has a rich heritage of launch vehicles that have used aerodynamic surfaces for flight stability such as the Saturn vehicles and flight control such as on the Redstone. Recently, due to aft center-of-gravity locations on launch vehicles currently being studied, the need has arisen for the vehicle control augmentation that is provided by these flight controls. Aerodynamic flight control can also reduce engine gimbaling requirements, provide actuator failure protection, enhance crew safety, and increase vehicle reliability, and payload capability. In the Saturn era, NASA went to the Moon with 300 sq ft of aerodynamic surfaces on the Saturn V. Since those days, the wealth of smart materials and advanced composites that have been developed allow for the design of very lightweight, strong, and innovative launch vehicle flight control surfaces. This paper presents an overview of the advanced composites and smart materials that are directly applicable to launch vehicle control surfaces.

Barret, C.↗
NASA NTRS2002

Natural Atmospheric Environment Model Development for the National Aeronautics and Space Administration's Second Generation Reusable Launch Vehicle

The National Aeronautics and Space Administration (NASA) recently began development of a new reusable launch vehicle. The program office is located at Marshall Space Flight Center (MSFC) and is called the Second Generation Reusable Launch Vehicle (2GRLV). The purpose of the program is to improve upon the safety and reliability of the first generation reusable launch vehicle, the Space Shuttle. Specifically, the goals are to reduce the risk of crew loss to less than 1-in-10,000 missions and decreased costs by a factor of 10 to approximately $1,000 per pound of payload launched to low Earth orbit. The program is currently in the very early stages of development and many two-stage vehicle concepts will be evaluated. Risk reduction activities are also taking place. These activities include developing new technologies and advancing current technologies to be used by the vehicle. The Environments Group at MSFC is tasked by the 2GRLV Program to develop and maintain an extensive series of analytical tools and environmental databases which enable it to provide detailed atmospheric studies in support of structural, guidance, navigation and control, and operation of the 2GRLV.

Roberts, Barry C.↗
NASA NTRS2018

A Prognostic Launch Vehicle Probability of Failure Assessment Methodology for Conceptual Systems Predicated on Human Causal Factors

Create an improved method to calculate reliability of a conceptual launch vehicle system prior to fabrication by using historic data of actual root causes of failures. While failures have unique "proximate causes", there are typically a finite amount of common "root causes". Heretofore launch vehicle reliability evaluation typically hardware-centric statistical analyses, while most root causes of failures are been shown to be human-centric. A method based on human-centric root causes can be used to quantify reliability assessments and focus proposed actions to mitigate problems. Existing methods have been optimistic in their projections of launch vehicle reliability compared to actuals. Hypothesis: reliability of a conceptual launch vehicle can be more accurately evaluated based on a rational, probabilistic approach using past failure assessment teams' findings predicated on human-centric causes."Human Reliability Analysis Methods Selection Guidance for NASA"Chandler F.T., et al., NASA HQ/OSMA study group, July 2006. Outside HRA experts from academia, other federal labs, and the private sector. 50 system reliability methods considered, fourteen selected for further study, four finally selected as best suited for human spaceflight. Probabilistic Risk Analysis (PRA) + Human Reliability Analysis (HRA) enabled incorporating effects and probabilities of human errors. While four down-selected methods deemed appropriate for failure assessment, it did not appear that these methods could be concisely applied to perform major system-wide assessment of probability of failure of a conceptual design without becoming unwieldy."Engineering a Safer World", Detailed, comprehensive study external to NASA Leveson N. G., MIT, 2011.Systems-Theoretic Accident Model and Processes (STAMP). All-encompassing accident model based on systems theory analyzed accidents after they occurred and created approaches to prevent occurrence in developing systems not focused on failure prevention per se, but rather reducing hazards by influencing human behavior through use of constraints, hierarchical control structures, and process models to improve system safetySystem Theoretic Process Analysis (STPA) addresses predictive part of problem (a "hazard analysis"). Includes all causal factors identified in STAMP: "...design errors, software flaws, component interaction accidents, cognitively complex human decision-making errors, and social organizational and management factors contributing to accidents" can guide design process rather than require it to exist before-hand did not appear capable of concise application for system-wide assessment of probability of failure of a conceptual design without becoming unwieldy.

Williams, Craig H.↗
NASA NTRS1971

Launch vehicle system requirements and restraints for the ERTS-A spacecraft

The technical requirements and restraints imposed by the ERTS spacecraft upon the Delta launch vehicle, shroud system, associated launch complex, and range are presented for technical coordination among various agencies involved in the launch vehicle and launch operations. The payload and spacecraft systems are described, and the mission, design, test, and launch base data are outlined.

Corrigan, J. F.↗
NASA NTRS1999

Launch Vehicle Design Process Description and Training Formulation

A primary NASA priority is to reduce the cost and improve the effectiveness of launching payloads into space. As a consequence, significant improvements are being sought in the effectiveness, cost, and schedule of the launch vehicle design process. In order to provide a basis for understanding and improving the current design process, a model has been developed for this complex, interactive process, as reported in the references. This model requires further expansion in some specific design functions. Also, a training course for less-experienced engineers is needed to provide understanding of the process, to provide guidance for its effective implementation, and to provide a basis for major improvements in launch vehicle design process technology. The objective of this activity is to expand the description of the design process to include all pertinent design functions, and to develop a detailed outline of a training course on the design process for launch vehicles for use in educating engineers whose experience with the process has been minimal. Building on a previously-developed partial design process description, parallel sections have been written for the Avionics Design Function, the Materials Design Function, and the Manufacturing Design Function. Upon inclusion of these results, the total process description will be released as a NASA TP. The design function sections herein include descriptions of the design function responsibilities, interfaces, interactive processes, decisions (gates), and tasks. Associated figures include design function planes, gates, and tasks, along with other pertinent graphics. Also included is an expanded discussion of how the design process is divided, or compartmentalized, into manageable parts to achieve efficient and effective design. A detailed outline for an intensive two-day course on the launch vehicle design process has been developed herein, and is available for further expansion. The course is in an interactive lecture/workshop format to engage the participants in active learning. The course addresses the breadth and depth of the process, requirements, phases, participants, multidisciplinary aspects, tasks, critical elements,as well as providing guidance from previous lessons learned. The participants are led to develop their own understanding of the current process and how it can be improved. Included are course objectives and a session-by-session outline of course content. Also included is an initial identification of visual aid requirements.

Atherton, James↗
NASA NTRS2011

Design Considerations for a Launch Vehicle Development Flight Instrumentation System

When embarking into the design of a new launch vehicle, engineering models of expected vehicle performance are always generated. While many models are well established and understood, some models contain design features that are only marginally known. Unfortunately, these analytical models produce uncertainties in design margins. The best way to answer these analytical issues is with vehicle level testing. The National Aeronautics and Space Administration respond to these uncertainties by using a vehicle level system called the Development Flight Instrumentation, or DFI. This DFI system can be simple to implement, with only a few measurements, or it may be a sophisticated system with hundreds of measurement and video, without a recording capability. From experience with DFI systems, DFI never goes away. The system is renamed and allowed to continue, in most cases. Proper system design can aid the transition to future data requirements. This paper will discuss design features that need to be considered when developing a DFI system for a launch vehicle. It will briefly review the data acquisition units, sensors, multiplexers and recorders, telemetry components and harnessing. It will present a reasonable set of requirements which should be implemented in the beginning of the program in order to start the design. It will discuss a simplistic DFI architecture that could be the basis for the next NASA launch vehicle. This will be followed by a discussion of the "experiences gained" from a past DFI system implementation, such as the very successful Ares I-X test flight. Application of these design considerations may not work for every situation, but they may direct a path toward success or at least make one pause and ask the right questions.

Johnson, Martin L.↗
NASA NTRS2016

Buckling Design and Imperfection Sensitivity of Sandwich Composite Launch-Vehicle Shell Structures

Composite materials are increasingly being considered and used for launch-vehicle structures. For shell structures, such as interstages, skirts, and shrouds, honeycomb-core sandwich composites are often selected for their structural efficiency. Therefore, it is becoming increasingly important to understand the structural response, including buckling, of sandwich composite shell structures. Additionally, small geometric imperfections can significantly influence the buckling response, including considerably reducing the buckling load, of shell structures. Thus, both the response of the theoretically perfect structure and the buckling imperfection sensitivity must be considered during the design of such structures. To address the latter, empirically derived design factors, called buckling knockdown factors (KDFs), were developed by NASA in the 1960s to account for this buckling imperfection sensitivity during design. However, most of the test-article designs used in the development of these recommendations are not relevant to modern launch-vehicle constructions and material systems, and in particular, no composite test articles were considered. Herein, a two-part study on composite sandwich shells to (1) examine the relationship between the buckling knockdown factor and the areal mass of optimized designs, and (2) to interrogate the imperfection sensitivity of those optimized designs is presented. Four structures from recent NASA launch-vehicle development activities are considered. First, designs optimized for both strength and stability were generated for each of these structures using design optimization software and a range of buckling knockdown factors; it was found that the designed areal masses varied by between 6.1% and 19.6% over knockdown factors ranging from 0.6 to 0.9. Next, the buckling imperfection sensitivity of the optimized designs is explored using nonlinear finite-element analysis and the as-measured shape of a large-scale composite cylindrical shell. When compared with the current buckling design recommendations, the results suggest that the current recommendations are overly conservative and that the development of new recommendations could reduce the acreage areal mass of many composite sandwich shell designs by between 4% and 19%, depending on the structure.

Schultz, Marc R.↗
NASA NTRS1996

Launch vehicle and power level impacts on electric GEO insertion

Solar Electric Propulsion (SEP) has been shown to increase net geosynchronous spacecraft mass when used for station keeping and final orbit insertion. The impact of launch vehicle selection and power level on the benefits of this approach were examined for 20 and 25 kW systems launched using the Ariane 5, Atlas IIAR, Long March, Proton, and Sea Launch vehicles. Two advanced on-board propulsion technologies, 5 kW ion and Hall thruster systems, were used to establish the relative merits of the technologies and launch vehicles. GaAs solar arrays were assumed. The analysis identifies the optimal starting orbits for the SEP orbit raising/plane changing while considering the impacts of radiation degradation in the Van Allen belts, shading, power degradation, and oblateness. This use of SEP to provide part of the orbit insertion results in net mass increases of 15 - 38% and 18 - 46% for one to two month trip times, respectively, over just using SEP for 15 years of north/south station keeping. SEP technology was shown to have a greater impact on net masses of launch vehicles with higher launch latitudes when avoidance of solar array and payload degradation is desired. This greater impact of SEP could help reduce the plane changing disadvantage of high latitude launch sites. Comparison with results for 10 and 15 kW systems show clear benefits of incremental increases in SEP power level, suggesting that an evolutionary approach to high power SEP for geosynchronous spacecraft is possible.

Oleson, Steven R.↗
NASA NTRS1977

Rain erosion considerations for launch vehicle insulation systems

In recent years the Delta launch vehicle has incorporated the capability to be launched through rain. This capability was developed to eliminate a design constraint which could result in a costly launch delay. This paper presents the methodology developed to implement rain erosion protection for the insulated exterior vehicle surfaces. The effect of the interaction between insulation material rain erosion resistance, rainstorm models, surface geometry and trajectory variations is examined. It is concluded that rain erosion can significantly impact the performance of launch vehicle insulation systems and should be considered in their design.

Daniels, D. J.↗