Analysis of space vehicle structures using the transfer-function concept.
Frequency transfer function applied to space vehicle structural analysis
Heer, E.↗
SEARCH · Engineering Papers
Results for “vehicle analysis”
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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Fundamentals of spacecraft thermal design.
The papers deal with and are grouped according to four major issues. These are: surface radiation properties, including synthesis and measurement, space flight effect, and contamination effects; thermal analysis, including reentry vehicle analysis, radiant heat transfer between surfaces, and thermal contact conductance of surfaces; heat pipes, including possible applications, operating characteristics, and design, fabrication and testing of heat pipes; and thermal design, including radiative, ablative, and active cooling thermal protection of the leading edge of a space-shuttle wing, and space station environmental thermal control. Individual items were previously announced in issues 06 and 11, 1971.
Lucas, J. W.↗
Parametric Systems Analysis of a Mid-Lift/Drag Entry System for Human Mars Mission
Mars entry systems are complex with many competing design objectives. This paper presents a data-driven and physics-based approach to systems analysis and tradespace exploration for a human Mars entry system using a Mid-Lift/Drag aerodynamic decelerator. The baseline entry system includes elements for Mars aerocapture, entry, descent, and landing segments of the mission. The approach uses a set of parametric models that enable a system-level analysis, vehicle sizing, sensitivity analysis, and tradespace exploration. The paper provides a brief introduction of the Mid-Lift/Drag concept, an overview of parametric models, discussions on alternative structural concepts, results from system-level analysis, and summary remarks.
Jamshid A Samareh↗
Hardware Verification and Validation for a Navigation Sensor Software Model in Support of Flight Vehicle Performance Analysis
… or, “It’s in the details, how to make complicated software perform like complicated hardware.” In attempts to minimize development time and quickly build an operational vehicle, NASA’s Space Launch System (SLS) has had to be intentional about integrated testing. Constraints on budget and schedule have required balance between testing needs and the desire for an integrated flight vehicle as soon as possible. To provide key insights early in design and analysis cycles, a large amount of effort has shifted into maturing and validating models at the component level with integrated testing as a means to validate their integration. In terms of SLS Navigation, this, and the model-based design approach have pushed explicit requirements for sensor models to be validated against flight hardware to high precision. This paper covers the approach taken to verify and validate the models for the two key navigation sensors on the SLS vehicle, the Redundant Inertial Navigation Sensor and the Rate Gyro Assembly. These models are used in performance evaluation, fault detection, and operations development extensively. Using a mix of data from hardware vendor documentation and testing reports, limited in-house testing, and integration activities, these models were able to be validated against flight hardware at multiple levels, from the internal software design to statistical behavior at the raw sensor and integrated box levels. The high level of insight into the hardware elements is instrumental to support flight certification activities and building confidence in SLS Navigation capability. Focused testing enabled additional insight and validation that proved invaluable and the resulting insights were used to focus and mature models. Additionally, of having validated performance-based hardware models enables a wide breadth of activities including detailed fault detection studies and integration into future vehicle frameworks, such as an upper stage and provide a valuable asset to continued SLS analysis and design.
Evan J Anzalone↗
Hardware Verification and Validation for a Navigation Sensor Software Model in Support of Flight Vehicle Performance Analysis
… or, “It’s in the details, how to make complicated software perform like complicated hardware.” In attempts to minimize development time and quickly build an operational vehicle, NASA’s Space Launch System (SLS) has had to be intentional about integrated testing. Constraints on budget and schedule have required balance between testing needs and the desire for an integrated flight vehicle as soon as possible. To provide key insights early in design and analysis cycles, a large amount of effort has shifted into maturing and validating models at the component level with integrated testing as a means to validate their integration. In terms of SLS Navigation, this, and the model-based design approach have pushed explicit requirements for sensor models to be validated against flight hardware to high precision. This paper covers the approach taken to verify and validate the models for the two key navigation sensors on the SLS vehicle, the Redundant Inertial Navigation Sensor and the Rate Gyro Assembly. These models are used in performance evaluation, fault detection, and operations development extensively. Using a mix of data from hardware vendor documentation and testing reports, limited in-house testing, and integration activities, these models were able to be validated against flight hardware at multiple levels, from the internal software design to statistical behavior at the raw sensor and integrated box levels. The high level of insight into the hardware elements is instrumental to support flight certification activities and building confidence in SLS Navigation capability. Focused testing enabled additional insight and validation that proved invaluable and the resulting insights were used to focus and mature models. Additionally, of having validated performance-based hardware models enables a wide breadth of activities including detailed fault detection studies and integration into future vehicle frameworks, such as an upper stage and provide a valuable asset to continued SLS analysis and design.
Thomas Park↗
Application of the V-Gamma map to vehicle breakup analysis
The V-Gamma map consists of all possible pairs of speed and flight path angle at atmospheric entry interface for accidental Earth reentries resulting from steady misaligned burns, incomplete burns, or no burn.
vehicle breakup analysis↗
Graphical analysis of Mars vehicle assembly
An analysis was performed to support the production of a computer graphics videotape of the assembly of an interplanetary vehicle at a man-tended transportation node, co-orbiting with Space Station Freedom. Using data provided by the Transportation Node Integration Agent of the Lunar-Mars Exploration Office, a workable design of the transportation node was developed. Kinematic manipulator simulation tools were used to perform the operations necessary to complete the assembly task. Results of this analysis are documented in the aforementioned video.
Lewis, Keven W.↗
Launch vehicle systems design analysis
Current launch vehicle design emphasis is on low life-cycle cost. This paper applies total quality management (TQM) principles to a conventional systems design analysis process to provide low-cost, high-reliability designs. Suggested TQM techniques include Steward's systems information flow matrix method, quality leverage principle, quality through robustness and function deployment, Pareto's principle, Pugh's selection and enhancement criteria, and other design process procedures. TQM quality performance at least-cost can be realized through competent concurrent engineering teams and brilliance of their technical leadership.
Ryan, Robert↗
Ares I-X First Flight Loss of Vehicle Probability Analysis
As part of the Constellation (Cx) Program development effort, several test flights were planned to prove concepts and operational capabilities of the new vehicles being developed. The first test, involving the Eastern Test Range, is the Ares I-X launched in 2009. As part of this test, the risk to the general public was addressed to ensure it is within Air Force requirements. This paper describes the methodology used to develop first flight estimates of overall loss of vehicle (LOV) failure probability, specifically for the Ares I-X. The method described in this report starts with the Air Force s generic failure probability estimate for first flight and adjusts the value based on the complexity of the vehicle as compared to the complexity of a generic vehicle. The results estimate a 1 in 9 probability of failure. The paper also describes traditional PRA methods used in this assessment, which were then combined with the updated first flight risk methodology to generate inputs required by the malfunction turn analysis to support estimate of casualty (Ec) calculations as part of the Final Flight Data Package (FFDP) delivered to the Eastern Range for Final Flight Plan Approval.
Bigler, Mark A.↗
Orion Launch Abort Vehicle Separation Analysis Using OVERFLOW
This slide presentation reviews the use of OVERFLOW, a flow solver, to analyze the effect of separation for a launch abort vehicle (i.e., Orion capsule) if required. Included in the presentation are views of the geometry, and the Overset grids, listing of the assumptions, the general run strategy, inputs into the Overflow solver, the required computational resources, the results of the convergence study. Charts and graphics are presented to show the results.
Booth, Tom↗
Launch Vehicle Systems Analysis
This report summaries the key accomplishments of Georgia Tech's Space Systems Design Laboratory (SSDL) under NASA Grant NAG8-1302 from NASA - Marshall Space Flight Center. The report consists of this summary white paper, copies of technical papers written under this grant, and several viewgraph-style presentations. During the course of this grant four main tasks were completed: (1)Simulated Combined-Cycle Rocket Engine Analysis Module (SCCREAM), a computer analysis tool for predicting the performance of various RBCC engine configurations; (2) Hyperion, a single stage to orbit vehicle capable of delivering 25,000 pound payloads to the International Space Station Orbit; (3) Bantam-X Support - a small payload mission; (4) International Trajectory Support for interplanetary human Mars missions.
Olds, John R.↗
Plasma Vehicle Charging Analysis for Orion Flight Test 1
In preparation for the upcoming experimental test flight for the Orion crew module, considerable interest was raised over the possibility of exposure to elevated levels of plasma activity and vehicle charging both externally on surfaces and internally on dielectrics during the flight test orbital operations. Initial analysis using NASCAP-2K indicated very high levels of exposure, and this generated additional interest in refining/defining the plasma and spacecraft models used in the analysis. This refinement was pursued, resulting in the use of specific AE8 and AP8 models, rather than SCATHA models, as well as consideration of flight trajectory, time duration, and other parameters possibly affecting the levels of exposure and the magnitude of charge deposition. Analysis using these refined models strongly indicated that, for flight test operations, no special surface coatings were necessary for the thermal protection system, but would definitely be required for future GEO, trans-lunar, and extra-lunar missions...
Lallement, L.↗
Plasma Vehicle Charging Analysis for Orion Flight Test 1
In preparation for the upcoming experimental test flight for the Orion crew module, considerable interest was raised over the possibility of exposure to elevated levels of plasma activity and vehicle charging both externally on surfaces and internally on dielectrics during the flight test orbital operations. Initial analysis using NASCAP-2K indicated very high levels of exposure, and this generated additional interest in refining/defining the plasma and spacecraft models used in the analysis. This refinement was pursued, resulting in the use of specific AE8 and AP8 models, rather than SCATHA models, as well as consideration of flight trajectory, time duration, and other parameters possibly affecting the levels of exposure and the magnitude of charge deposition. Analysis using these refined models strongly indicated that, for flight test operations, no special surface coatings were necessary for the Thermal Protection System (TPS), but would definitely be required for future GEO, trans-lunar, and extra-lunar missions.
Scully, B.↗
A fuselage/tank structure study for actively cooled hypersonic cruise vehicles: Structural analysis
The effects of fuselage cross-section (circular and elliptical) and structural arrangement (integral and nonintegral tanks) on the performance of actively cooled hypersonic cruise vehicles was evaluated. It was found that integrally machined stiffening of the tank walls, while providing the most weight-efficient use of materials, results in higher production costs. Fatigue and fracture mechanics appeared to have little effect on the weight of the three study aircraft. The need for thermal strain relief through insulation is discussed. Aircraft size and magnitude of the internal pressure are seen to be significant factors in tank design.
Baker, A. H.↗
Mars Sample Return Mars Ascent Vehicle Separation Analysis Utilizing the CLVTOPS Toolchain
A key element of the joint NASA and European Space Agency (ESA) Mars Sample Return (MSR) Campaign is the Mars Ascent Vehicle (MAV), which is being developed primarily by NASA Marshall Space Flight Center (MSFC), in association with NASA’s Jet Propulsion Laboratory (JPL) and Langley Research Center (LaRC). The MAV is a Mars-launched rocket that is responsible for transporting soil samples collected by the Perseverance rover from the Martian surface into orbit, where they will be captured by ESA's Earth Return Orbiter (ERO) for the return journey to Earth. The MAV design concept developed during the MAV Systems Requirement Cycle (SRC) and Preliminary Design Cycle (PDC) consisted of a two-solid-stage configuration, where the second stage is completely unguided in order to reduce vehicle and overall mission mass. The unguided second stage design presents technical challenges for the stage separation event, as the second stage trajectory and the payload’s ability to rendezvous with the ERO is extremely sensitive to disturbances during vehicle staging. The MSFC-developed CLVTOPS multibody dynamics toolchain was utilized to quickly assess multiple stage separation hardware options and to optimize the separation Concept of Operations (ConOps) in order to ensure successful near-field stage separation performance and maximize the orbital accuracy of the payload. This paper will describe how the CLVTOPS toolchain was used to assess the MAV stage separation event and help inform and optimize the MAV design and ConOps.
MSR↗
Structural Weight Estimation for Launch Vehicles
This paper describes some of the work in progress to develop automated structural weight estimation procedures within the Vehicle Analysis Branch (VAB) of the NASA Langley Research Center. One task of the VAB is to perform system studies at the conceptual and early preliminary design stages on launch vehicles and in-space transportation systems. Some examples of these studies for Earth to Orbit (ETO) systems are the Future Space Transportation System [1], Orbit On Demand Vehicle [2], Venture Star [3], and the Personnel Rescue Vehicle[4]. Structural weight calculation for launch vehicle studies can exist on several levels of fidelity. Typically historically based weight equations are used in a vehicle sizing program. Many of the studies in the vehicle analysis branch have been enhanced in terms of structural weight fraction prediction by utilizing some level of off-line structural analysis to incorporate material property, load intensity, and configuration effects which may not be captured by the historical weight equations. Modification of Mass Estimating Relationships (MER's) to assess design and technology impacts on vehicle performance are necessary to prioritize design and technology development decisions. Modern CAD/CAE software, ever increasing computational power and platform independent computer programming languages such as JAVA provide new means to create greater depth of analysis tools which can be included into the conceptual design phase of launch vehicle development. Commercial framework computing environments provide easy to program techniques which coordinate and implement the flow of data in a distributed heterogeneous computing environment. It is the intent of this paper to present a process in development at NASA LaRC for enhanced structural weight estimation using this state of the art computational power.
Cerro, Jeff↗
Launch Vehicle Loads Analysis Using Pressure-Sensitive Paint
Pressure transducers have been the instrumentation of choice for measuring unsteady flow phenomena. With recent advances in high-speed cameras, high-powered LEDs, and fast-response, pressure-sensitive paint, the unsteady pressure-sensitive paint (uPSP) technique has become a valuable alternative for production wind tunnel facilities, enabling time-resolved measurements of unsteady pressure fluctuations over a dense spatial grid on a wind tunnel model. Launch vehicle ground tests have proven to be a particularly well-matched application for uPSP due to the high signal level relative to tunnel background acoustics, relatively simple camera optical access, and rigidity of the model in wind-on test conditions. This presentation will highlight recent advances in data reduction of uPSP measurement data from recent launch vehicle wind tunnel tests at the NASA Ames Unitary Plan Wind Tunnel Complex (UPWT). The system can provide both localized surface pressure spectra as well as regional or zonal estimates of turbulence correlation model parameters. In addition, integrated vehicle-scale loads can be provided for buffet analysis. Data is reduced at the on-premise NASA Advanced Supercomputer (NAS) Division for just-in-time delivery of results during an ongoing wind tunnel test.
Pressure-Sensitive Paint↗
Launch Vehicle Loads Analysis Using Pressure-Sensitive Paint
Pressure transducers have been the instrumentation of choice for measuring unsteady flow phenomena. With recent advances in high-speed cameras, high-powered LEDs, and fast-response, pressure-sensitive paint, the unsteady pressure-sensitive paint (uPSP) technique has become a valuable alternative for production wind tunnel facilities, enabling time-resolved measurements of unsteady pressure fluctuations over a dense spatial grid on a wind tunnel model. Launch vehicle ground tests have proven to be a particularly well-matched application for uPSP due to the high signal level relative to tunnel background acoustics, relatively simple camera optical access, and rigidity of the model in wind-on test conditions. This presentation will highlight recent advances in data reduction of uPSP measurement data from recent launch vehicle wind tunnel tests at the NASA Ames Unitary Plan Wind Tunnel Complex (UPWT). The system can provide both localized surface pressure spectra as well as regional or zonal estimates of turbulence correlation model parameters. In addition, integrated vehicle-scale loads can be provided for buffet analysis. Data is reduced at the on-premise NASA Advanced Supercomputer (NAS) Division for just-in-time delivery of results during an ongoing wind tunnel test.
pressure-sensitive paint↗