SEARCH · Engineering Papers
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.
Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.
The flight experiment and operations plans for the Doppler Lidar System (DLS) are provided. Application of DLS to the study of severe storms and local weather penomena is addressed. Test plans involve 66 hours of flight time. Plans also include ground based severe storm and local weather data acquisition.
This paper describes the Dynamic Mode Decomposition (DMD) of the pressures on the scale model of the Space Launch System (SLS) Block 1 cargo vehicle with the Unsteady Pressure-Sensitive Paint (uPSP) measurements, which were collected in the Ascent Transient Aerodynamics Tests with the Unitary Plan Wind Tunnel 11-by-11-foot Transonic Wind Tunnel in September 2019 at NASA Ames Research Center. The work described in this paper is a part of NASA’s development of a new state-of-the-art uPSP capability in production wind tunnels. The conventional DMD algorithm is based on the Singular Value Decomposition (SVD) of the data matrix. For the matrix of the uPSP measurements of the SLS ATAT, the number of rows is equal to the number of nodes in the grid of the scale model, and the number of columns is equal to the number of frames in the videos taken with 4 Phantom high-speed cameras. In this paper, it is verified that, for the time series with zero mean value, the DMD is equivalent to the decomposition with the Discrete Fourier Transform (DFT). Considering the uPSP is mainly used in the assessment of the unsteady, aerodynamic phenomena, the DMD of the uPSP measurements can be implemented in two steps: (1) subtract the mean value from the uPSP measurement on each of the grid nodes; (2) apply the Fast Fourier Transform (FFT) on the resulting zero-mean time series. The DMD of the uPSP measurements with FFT has two advantages: (1) the computational complexity of FFT is O(N*logN), where N is the length of the time series; (2) compared to the SVD-based DMD algorithm, the DMD with FFT can be easily implemented in parallel processing. A sample matrix of uPSP measurements, at the size of 341 grid nodes and 128 frames, is generated. Figures 1 and 2 show the eigenvalues and the ratios of the eigenvectors, respectively, of the sample matrix, without and with the mean value removed on each of the grid nodes, computed with the SVD-based DMD and the FFT. The figures demonstrate the equivalence of the SVD-based DMD and the decomposition with DFT/FFT for the time series with zero mean value. The results of DMD of the uPSP measurements of the SLS ATAT in September 2019 are presented in the paper. The DMD modes at different frequencies are shown, the aerodynamic phenomena (e.g. shockwave and vortex shedding) are demonstrated and the correlation of the DMD modes with the test configuration parameter (e.g., the Mach Number) is discussed. Figure 3 shows a software tool to visualize the DMD modes. The code to implement the algorithm described in this paper was written in C, with libraries of FFTW for FFT and MPI/OpenMP for parallel processing, and executed on the NASA Pleiades supercomputer. Funding for this research was provided by the NASA Aerosciences Evaluation and Test Capabilities Project.
This paper discusses the thermal vacuum testing of a Self-Contained Heat Rejection Module (SHRM) currently being developed for use with Shuttle payloads and future spacecraft. Two weeks of testing were conducted in the NASA Johnson Space Center Space Environment Simulator Laboratory in October and November of 1975. The testing included demonstration of deployable spacecraft radiators with fluid swivels, variable return temperatures, and a dual-mode pumped fluid radiator/vapor compression refrigeration system. In addition, an evaluation was made of a contact heat exchanger used as the intercooler. The test articles and systems are described and the test planning and objectives discussed. A summary of the significant results of the testing and conclusions drawn from the testing are presented.
One of the key components of the Advanced Launch System (ALS) is a large liquid rocket, booster engine. To keep the overall vehicle size and cost down, this engine will probably use liquid oxygen (LOX) and a heavy hydrocarbon, such as RP-1, as propellants and operate at relatively high chamber pressures to increase overall performance. A technology program (Heavy Hydrocarbon Main Injector Technology) is being studied. The main objective of this effort is to develop a logic plan and supporting experimental data base to reduce the risk of developing a large scale (approximately 750,000 lb thrust), high performance main injector system. The overall approach and program plan, from initial analyses to large scale, two dimensional combustor design and test, and the current status of the program are discussed. Progress includes performance and stability analyses, cold flow tests of injector model, design and fabrication of subscale injectors and calorimeter combustors for performance, heat transfer, and dynamic stability tests, and preparation of hot fire test plans. Related, current, high pressure, LOX/RP-1 injector technology efforts are also briefly discussed.
Testgen is a computer program that generates suites of input and configuration vectors for testing other software or software/hardware systems. As systems become ever more complex, often, there is not enough time to test systems against all possible combinations of inputs and configurations, so test engineers need to be selective in formulating test plans. Testgen helps to satisfy this need: In response to a test-suite-requirement-specification model, it generates a minimal set of test vectors that satisfies all the requirements.
Based on the Apollo-era J-2 that powered the second and third stages of the Saturn V, the current J-2X is the liquid hydrogen and oxygen high-altitude rocket engine in development for both the Ares I Upper Stage and Ares V Earth Departure Stage. During my summer 2009 internship, J-2X was at a stage in its design maturity where verification testing needed to be considered for the benefit of adequate test facility preparation. My task was to focus on gimbal requirements and gimbal related hot-fire test plans. Facility capabilities were also of interest, specifically for hot-fire testing slated to occur at test stands A-1, A-2, and A-3 at Stennis Space Center(SSC) in Bay St. Louis, Mississippi. Gimbal requirements and stage interface conditions were investigated by applying a top-to-bottom systems engineering approach, which involved system level requirements, engine level requirements from both government and engine contractor perspectives, component level requirements, and the J-2X to Upper Stage and Earth Departure Stage interface control documents. Previous hydrogen and oxygen liquid rocket engine gimbal verification methods were researched for a glimpse at lessons learned. Discussion among the J-2X community affected by gimballing was organized to obtain input relative to proper verification of their respective component. Implementing suggestions such as gimbal pattern, angulated dwell time, altitude testing options, power level, and feed line orientation, I was able to match tests to test stands in the A Complex at SSC. Potential test capability gaps and risks were identified and pursued. The culmination of all these efforts was to coordinate with SSC to define additional facility requirements for both the A-3 altitude test stand that is currently under construction and the A-1 sea level test stand which is being renovated
Electric Vertical Takeoff and Landing (eVTOL) vehicles have the potential to enable cost effective Urban Air Mobility (UAM) applications. These concepts may also pose several challenging handling and control problems, which must be addressed prior to safe and reliable urban operations. This paper investigates a simplified vehicle control concept that is designed to address some of these challenges for a conceptual Lift Plus Cruise vehicle designed by NASA’s Revolutionary Vertical Lift Technology (RVLT) project. The command and control architecture for this concept is presented along with preliminary findings. Initial results explore the vehicle performance in an approach to hover transition scenario, designed to explore the challenge of dissipating energy across all flight regimes. Operational concepts with varying aggressiveness are evaluated through changing glideslope and deceleration rates. Preliminary results show that the simplified control concept is effective over these operational conditions, with control strategies and envelope protection limits able to maintain control though aggressive operations despite saturation at steeper slopes with higher deceleration rates. Final results will show piloted simulation evaluations in the the Aerospace Cognitive Engineering Lab – Rapid Automation Test Environment (ACEL-RATE) laboratory at NASA Ames Research Center. The planned tests will build on these operations with additional test cases exploring variations in wind conditions as well as transition-to-hover automation strategies and display information.
A robust and experienced systems engineering team is essential for a successful program. It is often a challenge to build a core systems engineering team early enough in a program to maximize integration and assure a common path for all supporting teams in a project. Ares I was no exception. During the planning of IVGVT, the team had many challenges including lack of: early identification of stakeholders, team training in NASA s system engineering practices, solid requirements flow down and a top down documentation strategy. The IVGVT team started test planning early in the program before the systems engineering framework had been matured due to an aggressive schedule. Therefore the IVGVT team increased their involvement in the Constellation systems engineering effort. Program level requirements were established that flowed down to IVGVT aligning all stakeholders to a common set of goals. The IVGVT team utilized the APPEL REQ Development Management course providing the team a NASA focused model to follow. The IVGVT team engaged directly with the model verification and validation process to assure that a solid set of requirements drove the need for the test event. The IVGVT team looked at the initial planning state, analyzed the current state and then produced recommendations for the ideal future state of a wide range of systems engineering functions and processes. Based on this analysis, the IVGVT team was able to produce a set of lessons learned and to provide suggestions for future programs or tests to use in their initial planning phase.
A program designed to implement a ceramic matrix composite integrally bladed disk (blisk) into rocket engine style turbomachinery has successfully completed testing. The Marshall Space Flight Center (MSFC) program, utilizing the MSFC turbomachinery design, analysis, and testing capabilities along with materials development capabilities from both Glenn Research Center (GRC) and MSFC, has tested two carbon fiber reinforced silicon carbide blisks in the Simplex Turbopump at MSFC's Test Stand 500. One blisk contained a polar woven fiber preform, while the second blisk tested utilized a quasi-isotropic preform. Vhile earlier papers have chronicled the program's design, material testing, and torque testing efforts, this paper focuses on the testing of the blisks in the Simplex turbopump. Emphasis will be placed on the actual condition of the blisks before and after the testing test program design methodology, and conclusions that can be drawn from the test data and blisk final conditions. The program performed three separate test series. The first series was needed to validate that the Simplex turbopump was correctly re-built following a major incident to the turbopump. The turbopump had two major differences from the original design. The most obvious difference was the sleeve required throughout the bore of the main housing. The second major difference was modifications to the pump diffuser to improve performance. Several areas were burnt during the incident and were either repaired by weld repair (pump inlet housing) or simply smoothed out (turbine nozzle discharge). The test series was designed to weed out any turbopump design and manufacturing flaws or fatigue issues prior to putting the C/SiC blisks into it. The second and third series were the C/SiC blisk test series. The primary goal of these series was to expose the blisks to as much fatigue causing dynamic stress as possible to examine the material's capability. Initially, the test plan was to put equal time on the two blisks, however, as the test series progressed, the funding allowed additional testing to occur. The additional test time was placed on the polar weave blisk. The total test time accrued on the polar blisk was 2550 seconds with 860 seconds near the turbopump design speed of 25,000 rpm. This testing included 6 tests / 775 seconds pumping liquid nitrogen and 7 tests / 1775 seconds pumping liquid oxygen. The drive gas for all of the tests was gaseous nitrogen due to the lack of hot gas source for the Simplex turbopump. The quasi-isotropic blisk was tested for XX total tests and XXXX seconds with X tests/XXXX seconds pumping liquid nitrogen and X tests/XXXX seconds pumping liquid \oxygen.During the test series, the blisks were inspected following each test. Inspections initially were viewed from the downstream side of the blisks only. Midway through the testing, a method of borescoping the leading edges of the blades was devised, and subsequently, both sides of the blades were inspected following each test. The leading and trailing edges of the polar blisk held up better than the quasi-isotropic blisk. This was a known possibility due to the varying fiber direction in the blades as the rectangular preform weave is cut in a circular pattern. The surprising fact about the testing was that there was no measurable performance loss due to the inaccuracies in the blade manufacturing in the C/SiC blisks, the surface roughness C/SiC of the blades, or the loss of the material in the polar blisk. A performance shift was seen in the quasi-isotropic blisk as portions of the leading and trailing edges were lost. After the testing was completed, detailed inspections of the blisks were performed. The largest surprise was the polar blisk had a obvious crack in a single blade that was located nearly midspan which was not detected in test. The crack ran completely through the blade circumferenciary and through the radial length of the blade. However, the crack does not appear to extend into the blisk hub. Although the cause of the crack is still under investigation, the material appears to be tolerant of this crack, and other hairline cracks discovered under higher magnification. This bodies well for eventual use of this material in actual flight turbopumps where monolithic fracture toughness issues limit its use.
The National Aeronautics and Space Administration (NASA) Access 5 Project Office sponsored a cooperative collision avoidance flight demonstration program for unmanned aircraft systems (UAS). This flight test was accomplished between September 21st and September 27th 2005 from the Mojave Airport, Mojave, California. The objective of these flights was to collect data for the Access 5 Cooperative Collision Avoidance (CCA) Work Package simulation effort, i.e., to gather data under select conditions to allow validation of the CCA simulation. Subsequent simulation to be verified were: Demonstrate the ability to detect cooperative traffic and provide situational awareness to the ROA pilot; Demonstrate the ability to track the detected cooperative traffic and provide position information to the ROA pilot; Demonstrate the ability to determine collision potential with detected cooperative traffic and provide notification to the ROA pilot; Demonstrate that the CCA subsystem provides information in sufficient time for the ROA pilot to initiate an evasive maneuver to avoid collision; Demonstrate an evasive maneuver that avoids collision with the threat aircraft; and lastly, Demonstrate the ability to assess the adequacy of the maneuver and determine that the collision potential has been avoided. The Scaled Composites, LLC Proteus Optionally Piloted Vehicle (OPV) was chosen as the test platform. Proteus was manned by two on-board pilots but was also capable of being controlled from an Air Vehicle Control Station (AVCS) located on the ground. For this demonstration, Proteus was equipped with cooperative collision sensors and the required hardware and software to place the data on the downlink. Prior to the flight phase, a detailed set of flight test scenarios were developed to address the flight test objectives. Two cooperative collision avoidance sensors were utilized for detecting aircraft in the evaluation: Traffic Alert and Collision Avoidance System-II (TCAS-II) and Automatic Dependent Surveillance Broadcast (ADS-B). A single intruder aircraft was used during all the flight testing, a NASA Gulfstream III (G-III). During the course of the testing, six geometrically different near-collision scenarios were evaluated. These six scenarios were each tested using various combinations of sensors and collision avoidance software. Of the 54 planned test points 49 were accomplished successfully. Proteus flew a total of 21.5 hours during the testing and the G-III flew 19.8 hours. The testing fully achieved all flight test objectives. The Flight IPT performed an analysis to determine the accuracy of the simulation model used to predict the location of the host aircraft downstream during an avoidance maneuver. The data collected by this flight program was delivered to the Access 5 Cooperative Collision Avoidance (CCA) Work Package Team who was responsible for reporting on their analysis of this flight data.
This paper describes the approach used by the Space Communication and Navigation (SCaN) Testbed team to qualify three Software Defined Radios (SDR) for operation in space and the characterization of the platform to enable upgrades on-orbit. The three SDRs represent a significant portion of the new technologies being studied on board the SCAN Testbed, which is operating on an external truss on the International Space Station (ISS). The SCaN Testbed provides experimenters an opportunity to develop and demonstrate experimental waveforms and applications for communication, networking, and navigation concepts and advance the understanding of developing and operating SDRs in space. Qualifying a Software Defined Radio for the space environment requires additional consideration versus a hardware radio. Tests that incorporate characterization of the platform to provide information necessary for future waveforms, which might exercise extended capabilities of the hardware, are needed. The development life cycle for the radio follows the software development life cycle, where changes can be incorporated at various stages of development and test. It also enables flexibility to be added with minor additional effort. Although this provides tremendous advantages, managing the complexity inherent in a software implementation requires a testing beyond the traditional hardware radio test plan. Due to schedule and resource limitations and parallel development activities, the subsystem testing of the SDRs at the vendor sites was primarily limited to typical fixed transceiver type of testing. NASA s Glenn Research Center (GRC) was responsible for the integration and testing of the SDRs into the SCaN Testbed system and conducting the investigation of the SDR to advance the technology to be accepted by missions. This paper will describe the unique tests that were conducted at both the subsystem and system level, including environmental testing, and present results. For example, test waveforms were developed to measure the gain of the transmit system across the tunable frequency band. These were used during thermal vacuum testing to enable characterization of the integrated system in the wide operational temperature range of space. Receive power indicators were used for Electromagnetic Interference tests (EMI) to understand the platform s susceptibility to external interferers independent of the waveform. Additional approaches and lessons learned during the SCaN Testbed subsystem and system level testing will be discussed that may help future SDR integrators
This paper describes the approach used by the Space Communication and Navigation (SCaN) Testbed team to qualify three Software Defined Radios (SDR) for operation in space and the characterization of the platform to enable upgrades on-orbit. The three SDRs represent a significant portion of the new technologies being studied on board the SCAN Testbed, which is operating on an external truss on the International Space Station (ISS). The SCaN Testbed provides experimenters an opportunity to develop and demonstrate experimental waveforms and applications for communication, networking, and navigation concepts and advance the understanding of developing and operating SDRs in space. Qualifying a Software Defined Radio for the space environment requires additional consideration versus a hardware radio. Tests that incorporate characterization of the platform to provide information necessary for future waveforms, which might exercise extended capabilities of the hardware, are needed. The development life cycle for the radio follows the software development life cycle, where changes can be incorporated at various stages of development and test. It also enables flexibility to be added with minor additional effort. Although this provides tremendous advantages, managing the complexity inherent in a software implementation requires a testing beyond the traditional hardware radio test plan. Due to schedule and resource limitations and parallel development activities, the subsystem testing of the SDRs at the vendor sites was primarily limited to typical fixed transceiver type of testing. NASA's Glenn Research Center (GRC) was responsible for the integration and testing of the SDRs into the SCaN Testbed system and conducting the investigation of the SDR to advance the technology to be accepted by missions. This paper will describe the unique tests that were conducted at both the subsystem and system level, including environmental testing, and present results. For example, test waveforms were developed to measure the gain of the transmit system across the tunable frequency band. These were used during thermal vacuum testing to enable characterization of the integrated system in the wide operational temperature range of space. Receive power indicators were used for Electromagnetic Interference tests (EMI) to understand the platform's susceptibility to external interferers independent of the waveform. Additional approaches and lessons learned during the SCaN Testbed subsystem and system level testing will be discussed that may help future SDR integrators.
The International Space Station (ISS) is a result of international collaboration in building a sophisticated laboratory of an unprecedented scale in Low Earth Orbit. After a complex assembly sequence spanning over a decade, some of the early modules launched at the beginning of the program would reach the end of their certified lives, while the newer modules were just being commissioned into operation. To maximize the return on global investments in this one-of-a-kind orbiting platform that was initially conceived for a service life until 2016, it is essential for the cutting edge research on ISS to continue as long as the station can be sustained safely in orbit. ISS Program is assessing individual modules in detail to extend the service life of the ISS to 2024, and possibly to 2028. Without life extension, Functional Cargo Block (known by its Russian acronym as FGB) and the Service Module (SM), two of the early modules on the Russian Segment, would reach the end of their certified lives in 2013 and 2015 respectively. Both FGB and SM are critical for the propulsive function of the ISS. This paper describes the approach used for the service life extension of the FGB propulsion system. Also presented is an overview of the system description along with the process adopted for developing the life test plans based on considerations of system failure modes, fault tolerance and safety provisions. Tests and analyses performed, important findings and life estimates are summarized. Based on the life extension data, FGB propulsion system, in general, is considered ready for a service life until 2028.
The Armstrong Flight Research Center has performed loads testing of a series of developmental atmospheric entry decelerator structural components. Test setup hardware were designed and fabricated. In addition, test plan and checklist were developed for the consistent and efficient execution of the tests. Eight test articles were successfully tested in over one hundred test runs as test objectives were met. Test article buckling shapes and buckling loads were observed. Displacements and strains were also recorded as various load cases were applied. The test data was sent to Langley Research Center to help with the construction of the finite element model of the decelerator assembly.
This short presentation provides an overview of the technical work that has been performed, and is planned, under the Advanced Air Mobility project's Hazard Perception and Avoidance (HPA) technical area. The presentation starts with a brief overview of the HPA technical area. The brief then covers a recently-completed part task human-in-the-loop simulation and outlines a follow-on simulation that is planned for August. The next portion of the brief outlines HPA's flight test plans, including some details on the Integration of Automated Systems (IAS) flights planned for 2023. The presentation ends with a high level overview of the activies planned in the out years of the HPA project.
This short presentation provides an overview of the technical work that has been performed, and is planned, under the Advanced Air Mobility project's Hazard Perception and Avoidance (HPA) technical area. The presentation starts with a brief overview of the HPA technical area. The brief then covers a recently-completed part task human-in-the-loop simulation and outlines a follow-on simulation that is planned for August. The next portion of the brief outlines HPA's flight test plans, including some details on the Integration of Automated Systems (IAS) flights planned for 2023. The presentation ends with a high level overview of the activies planned in the out years of the HPA project.
A summary of a ten week project on flexible multibody modeling, verification and control is presented. Emphasis was on the need for experimental verification. A literature survey was conducted for gathering information on the existence of experimental work related to flexible multibody systems. The first portion of the assigned task encompassed the modeling aspects of flexible multibodies that can undergo large angular displacements. Research in the area of modeling aspects were also surveyed, with special attention given to the component mode approach. Resulting from this is a research plan on various modeling aspects to be investigated over the next year. The relationship between the large angular displacements, boundary conditions, mode selection, and system modes is of particular interest. The other portion of the assigned task was the generation of a test plan for experimental verification of analytical and/or computer analysis techniques used for flexible multibody systems. Based on current and expected frequency ranges of flexible multibody systems to be used in space applications, an initial test article was selected and designed. A preliminary TREETOPS computer analysis was run to ensure frequency content in the low frequency range, 0.1 to 50 Hz. The initial specifications of experimental measurement and instrumentation components were also generated. Resulting from this effort is the initial multi-phase plan for a Ground Test Facility of Flexible Multibody Systems for Modeling Verification and Control. The plan focusses on the Multibody Modeling and Verification (MMV) Laboratory. General requirements of the Unobtrusive Sensor and Effector (USE) and the Robot Enhancement (RE) laboratories were considered during the laboratory development.
The Thiokol, Science and Engineering Huntsville Operations (SEHO) Laboratory has previously experienced significant levels of variation in testing Quadruple Lap Shear (QLS) specimens. The QLS test is used at Thiokol / Utah for the qualification of Reusable Solid Rocket Motor (RSRM) nozzle flex bearing materials. A test was conducted to verify that process changes instituted by SEHO personnel effectively reduced variability, even with normal processing variables introduced. A test matrix was designed to progress in a series of steps; the first establishing a baseline, then introducing additional solvents or other variables. Variables included normal test plan delay times, pre-bond solvent hand-wipes and contaminants. Each condition tested utilized standard QLS hardware bonded with natural rubber, two separate technicians and three replicates. This paper will report the results and conclusions of this investigation.