Gemini launch vehicle guidance and performance
Flight test results of Gemini launch vehicle guidance system
FLIGHT TEST↗
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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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A computer simulation of Skylab dynamics and attitude control for performance verification and operational support
A simulation of the Skylab attitude and pointing control system (APCS) is outlined and discussed. Implementation is via a large hybrid computer and includes those factors affecting system momentum management, propellant consumption, and overall vehicle performance. The important features of the flight system are discussed; the mathematical models necessary for this treatment are outlined; and the decisions involved in implementation are discussed. A brief summary of the goals and capabilities of this tool is also included.
Buchanan, H.↗
Configuration evaluation and criteria plan. Volume 2: Evaluation criteria plan (update). Space Transportation Main Engine (STME) configuration study
Candidate main engine configurations which enhance vehicle performance, operation and cost are identified. These candidate configurations are evaluated and the configurations which provide significant advantages over existing systems are selected for consideration for the next generation of launch vehicles. The unbiased selection of the Space Transportation Main Engine (STME) configuration requires that the candidate engines be evaluated against a predetermined set of criteria which must be properly weighted to emphasize critical requirements defined prior to the actual evaluation. During a prior study of the STME a Gas Generator Cycle engine was selected for conceptual design, with emphasis on reusability, reliability and low cost while achieving good performance. In this study emphasis is on expendable application of the STME while maintaining low cost and high reliability.
Bair, E. K.↗
The QED engine - Fusion-electric propulsion for Cis-Oort/Quasi-Interstellar (QIS) flight
A summary is presented of QED fusion-direct-electric engine systems, their features, and performance ranges. The principles and characteristics of inertial-electrostatic-fusion (IEF) power source systems are then reviewed, and their application to the diluted-fusion-product (DFP) engine concept for QIS missions is discussed. Particular attention is given to vehicle performance over a range of very high specific impulses and to specifications of a typical candidate DFP/IEF engine and a single-stage vehicle for rapid flight to 550 AU.
Bussard, Robert W.↗
Aerosciences, Aero-Propulsion and Flight Mechanics Technology Development for NASA's Next Generation Launch Technology Program
The Next Generation Launch Technology (NGLT) program, Vehicle Systems Research and Technology (VSR&T) project is pursuing technology advancements in aerothermodynamics, aeropropulsion and flight mechanics to enable development of future reusable launch vehicle (RLV) systems. The current design trade space includes rocket-propelled, hypersonic airbreathing and hybrid systems in two-stage and single-stage configurations. Aerothermodynamics technologies include experimental and computational databases to evaluate stage separation of two-stage vehicles as well as computational and trajectory simulation tools for this problem. Additionally, advancements in high-fidelity computational tools and measurement techniques are being pursued along with the study of flow physics phenomena, such as boundary-layer transition. Aero-propulsion technology development includes scramjet flowpath development and integration, with a current emphasis on hypervelocity (Mach 10 and above) operation, as well as the study of aero-propulsive interactions and the impact on overall vehicle performance. Flight mechanics technology development is focused on advanced guidance, navigation and control (GN&C) algorithms and adaptive flight control systems for both rocket-propelled and airbreathing vehicles.
Cockrell, Charles E., Jr.↗
Analysis of the Tropospheric Doppler Radar Wind Profiler Measurement Accuracy
Space launch vehicle trajectory design uses atmospheric winds to determine vehicle performance and structural margin assessments prior to flight. For launches at the United States Space Force’s Eastern Range (ER) at the Cape Canaveral Space Force Station (CCSFS), tropospheric wind measurements come from either in-situ or remote sensing instrumentation. The use of the National Aeronautics and Space Administration’s (NASA) Kennedy Space Center (KSC) 48-MHz Tropospheric Doppler Radar Wind Profiler (TDRWP) has become the primary wind measurement source for several launch vehicles. Extensive evaluations of TDRWP data have occurred to quantify the system performance. This includes quantifying the accuracy of wind estimates over the entire sampling altitude as well as at discrete altitudes. This paper will describe methodology and analyses used to quantify the measurement accuracy of TDRWP data when operating in an alternative mission support mode over a 2-year period from March 2020 to March 2022.
Meghan E. Carrico↗
Analysis of the Tropospheric Doppler Radar Wind Profiler Measurement Accuracy
Space launch vehicle trajectory design uses atmospheric winds to determine vehicle performance and structural margin assessments prior to flight. For launches at the United States Space Force’s Eastern Range (ER) at the Cape Canaveral Space Force Station (CCSFS), tropospheric wind measurements come from either in-situ or remote sensing instrumentation. The use of the National Aeronautics and Space Administration’s (NASA) Kennedy Space Center (KSC) 48-MHz Tropospheric Doppler Radar Wind Profiler (TDRWP) has become the primary wind measurement source for several launch vehicles. Extensive evaluations of TDRWP data have occurred to quantify the system performance. This includes quantifying the accuracy of wind estimates over the entire sampling altitude as well as at discrete altitudes. This paper will describe methodology and analyses used to quantify the measurement accuracy of TDRWP data when operating in an alternative mission support mode over a 2-year period from March 2020 to March 2022.
Meghan E Carrico↗
Effects of stick dynamics on helicopter flying qualities
An experiment that investigated the influence of typical helicopter force-feel system dynamics on roll-axis handling qualities was conducted in concurrent ground and inflight simulations. Variations in lateral control natural frequency and damping ratio, effected by changes in inertia and damping, were evaluated in a disturbance-rejection task. Pilot ratings indicated a preference for low-inertia feel systems, although measured performance was relatively constant over the range of stick characteristics. Force-sensing was compared with position sensing as the input to the control system. Force-sensing improved performance but did not improve pilot ratings. Overall, the results indicated that control-stick dynamics, at least within a reasonable range, did not have a significant effect on pilot-vehicle performance. However, the physical effort required to maintain a desired pilot/manipulator bandwidth became objectionable as the stick inertia increased beyond 5-7 lbm, which was reflected in the pilot ratings and comments.
Watson, Douglas C.↗
Feasibility of Turing-Style Tests for Autonomous Aerial Vehicle "Intelligence"
A new approach is suggested to define and evaluate key metrics as to autonomous aerial vehicle performance. This approach entails the conceptual definition of a "Turing Test" for UAVs. Such a "UAV Turing test" would be conducted by means of mission simulations and/or tailored flight demonstrations of vehicles under the guidance of their autonomous system software. These autonomous vehicle mission simulations and flight demonstrations would also have to be benchmarked against missions "flown" with pilots/human-operators in the loop. In turn, scoring criteria for such testing could be based upon both quantitative mission success metrics (unique to each mission) and by turning to analog "handling quality" metrics similar to the well-known Cooper-Harper pilot ratings used for manned aircraft. Autonomous aerial vehicles would be considered to have successfully passed this "UAV Turing Test" if the aggregate mission success metrics and handling qualities for the autonomous aerial vehicle matched or exceeded the equivalent metrics for missions conducted with pilots/human-operators in the loop. Alternatively, an independent, knowledgeable observer could provide the "UAV Turing Test" ratings of whether a vehicle is autonomous or "piloted." This observer ideally would, in the more sophisticated mission simulations, also have the enhanced capability of being able to override the scripted mission scenario and instigate failure modes and change of flight profile/plans. If a majority of mission tasks are rated as "piloted" by the observer, when in reality the vehicle/simulation is fully- or semi- autonomously controlled, then the vehicle/simulation "passes" the "UAV Turing Test." In this regards, this second "UAV Turing Test" approach is more consistent with Turing s original "imitation game" proposal. The overall feasibility, and important considerations and limitations, of such an approach for judging/evaluating autonomous aerial vehicle "intelligence" will be discussed from a theoretical perspective.
Young, Larry A.↗
Three Orbital Burns to Molniya Orbit via Shuttle Centaur G Upper Stage
An unclassified analytical trajectory design, performance, and mission study was done for the 1982-86 joint NASA-USAF Shuttle/Centaur G upper stage development program to send performance-demanding payloads to high orbits such as Molniya using an unconventional orbit transfer. This optimized three orbital burn transfer to Molniya orbit was compared to the then-baselined two burn transfer. The results of the three dimensional trajectory optimization performed include powered phase steering data and coast phase orbital element data. Time derivatives of the orbital elements as functions of thrust components were evaluated and used to explain the optimization's solution. Vehicle performance as a function of parking orbit inclination was given. Performance and orbital element data was provided for launch windows as functions of launch time. Ground track data was given for all burns and coasts including variation within the launch window. It was found that a Centaur with fully loaded propellant tanks could be flown from a 37deg inclination low Earth parking orbit and achieve Molniya orbit with comparable performance to the baselined transfer which started from a 57deg inclined orbit: 9,545 lb vs. 9,552 lb of separated spacecraft weight respectively. There was a significant reduction in the need for propellant launch time reserve for a one hour window: only 78 lb for the three burn transfer vs. 320 lb for the two burn transfer. Conversely, this also meant that longer launch windows over more orbital revolutions could be done for the same amount of propellant reserve. There was no practical difference in ground tracking station or airborne assets needed to secure telemetric data, even though the geometric locations of the burns varied considerably. There was a significant adverse increase in total mission elapsed time for the three vs. two burn transfer (12 vs. 11/4 hrs), but could be accommodated by modest modifications to Centaur systems. Future applications were discussed. The three burn transfer was found to be a viable, arguably preferable, alternative to the two burn transfer.
Orbital trajectory design↗
Three Orbital Burns to Molniya Orbit Via Shuttle_Centaur G Upper Stage
An unclassified analytical trajectory design, performance, and mission study was done for the 1982 to 1986 joint National Aeronautics and Space Administration (NASA)-United States Air Force (USAF) Shuttle/Centaur G upper stage development program to send performance-demanding payloads to high orbits such as Molniya using an unconventional orbit transfer. This optimized three orbital burn transfer to Molniya orbit was compared to the then-baselined two burn transfer. The results of the three dimensional trajectory optimization performed include powered phase steering data and coast phase orbital element data. Time derivatives of the orbital elements as functions of thrust components were evaluated and used to explain the optimization's solution. Vehicle performance as a function of parking orbit inclination was given. Performance and orbital element data was provided for launch windows as functions of launch time. Ground track data was given for all burns and coasts including variation within the launch window. It was found that a Centaur with fully loaded propellant tanks could be flown from a 37 deg inclination low Earth parking orbit and achieve Molniya orbit with comparable performance to the baselined transfer which started from a 57 deg inclined orbit: 9,545 versus 9,552 lb of separated spacecraft weight, respectively. There was a significant reduction in the need for propellant launch time reserve for a 1 hr window: only 78 lb for the three burn transfer versus 320 lb for the two burn transfer. Conversely, this also meant that longer launch windows over more orbital revolutions could be done for the same amount of propellant reserve. There was no practical difference in ground tracking station or airborne assets needed to secure telemetric data, even though the geometric locations of the burns varied considerably. There was a significant adverse increase in total mission elapsed time for the three versus two burn transfer (12 vs. 1-1/4 hr), but could be accommodated by modest modifications to Centaur systems. Future applications were discussed. The three burn transfer was found to be a viable, arguably preferable, alternative to the two burn transfer.
Orbital trajectory design↗
Space Launch System Mobile Launcher Modal Pretest Analysis
NASA is developing an expendable heavy lift launch vehicle capability, the Space Launch System, to support lunar and deep space exploration. To support this capability, an updated ground infrastructure is required including modifying an existing Mobile Launcher system. The Mobile Launcher is a very large heavy beam/truss steel structure designed to support the Space Launch System during its buildup and integration in the Vehicle Assembly Building, transportation from the Vehicle Assembly Building out to the launch pad, and provides the launch platform at the launch pad. The previous Saturn/Apollo and Space Shuttle programs had integrated vehicle ground vibration tests of their integrated launch vehicles performed with simulated free-free boundary conditions to experimentally anchor and validate structural and flight controls analysis models. For the Space Launch System program, the Mobile Launcher will be used as the modal test fixture for the ground vibration test of the first Space Launch System flight vehicle, Exploration Mission ? 1( now referred to as Artemis 1), programmatically referred to as the Integrated vehicle modal test. The Integrated vehicle modal test of the Exploration Mission - 1 integrated launch vehicle will have its core and second stages unfueled while mounted to the ML while inside the Vehicle Assembly Building, which is currently scheduled for the late spring or early summer of 2020. The Space Launch System program has implemented a building block approach for dynamic model validation. The modal test of the Mobile Launcher is an important part of this building block approach in supporting the integrated vehicle modal test since the Mobile Launcher will serve as a structurally dynamic test fixture whose modes will couple with the modes of the Exploration Mission ? 1 test vehicle. The Mobile Launcher modal test will further support understanding the structural dynamics of the Mobile Launcher and SLS during rollout to the launch pad, which will play a key role in better understanding and prediction of the rollout forces acting on the launch vehicle. The Mobile Launcher modal test is currently scheduled for the summer of 2019. Due to a very tight modal testing schedule, this Mobile Launcher modal pretest analysis has been performed to ensure there is a high likelihood of being able to successfully complete the modal test (i.e. identify the primary target modes) using the planned instrumentation, shakers, and excitation types. This paper will discuss this Mobile Launcher modal pretest analysis and the unique challenges faced due to the Mobile Launcher's size and weight, which are typically not faced when modal testing aerospace structures.
Akers, James C.↗
Space Launch System Mobile Launcher Modal Pretest Analysis
NASA is developing an expendable heavy lift launch vehicle capability, the Space Launch System, to support lunar and deep space exploration. To support this capability, an updated ground infrastructure is required including modifying an existing Mobile Launcher system. The Mobile Launcher is a very large heavy beam/truss steel structure designed to support the Space Launch System during its buildup and integration in the Vehicle Assembly Building, transportation from the Vehicle Assembly Building out to the launch pad, and provides the launch platform at the launch pad. The previous Saturn/Apollo and Space Shuttle programs had integrated vehicle ground vibration tests of their integrated launch vehicles performed with simulated free-free boundary conditions to experimentally anchor and validate structural and flight controls analysis models. For the Space Launch System program, the Mobile Launcher will be used as the modal test fixture for the ground vibration test of the first Space Launch System flight vehicle, Artemis 1, programmatically referred to as the integrated vehicle modal test. The integrated vehicle modal test of the Artemis 1 integrated launch vehicle will have its core and second stages unfueled while mounted to the Mobile Launcher while inside the Vehicle Assembly Building, which is currently scheduled for the summer of 2020. The Space Launch System program has implemented a building block approach for dynamic model validation. The modal test of the Mobile Launcher is an important part of this building block approach in supporting the integrated vehicle modal test since the Mobile Launcher will serve as a structurally dynamic test fixture whose modes will couple with the modes of the Artemis 1 integrated vehicle. The Mobile Launcher modal test will further support understanding the structural dynamics of the Mobile Launcher and Space Launch System during rollout to the launch pad, which will play a key role in better understanding and prediction of the rollout forces acting on the launch vehicle. The Mobile Launcher modal test is currently scheduled for the summer of 2019. Due to a very tight modal testing schedule, this independent Mobile Launcher modal pretest analysis has been performed to ensure there is a high likelihood of successfully completing the modal test (i.e. identify the primary target modes) using the planned instrumentation, shakers, and excitation types. This paper will discuss this Mobile Launcher modal pretest analysis for its three test configurations and the unique challenges faced due to the Mobile Launcher’s size and weight, which are typically not faced when modal testing aerospace structures.
Akers, James C.↗
Evaluation of half wave induction motor drive for use in passenger vehicles
Research performed at the University of Missouri-Columbia to devise and design a lower cost inverter induction motor drive for electrical propulsion of passenger vehicles is described. A two phase inverter motor system is recommended. The new design is predicted to provide comparable vehicle performance, improved reliability and a cost advantage for a high production vehicle, decreased total rating of the power semiconductor switches, and a somewhat simpler control hardware compared to the conventional three phase bridge inverter motor drive system. The major disadvantages of the two phase inverter motor drive are that it is larger and more expensive than a three phase machine, the design of snubbers for the power leakage inductances produce higher transient voltages, and the torque pulsations are relatively large because of the necessity to limit the inverter switching frequency to achieve high efficiency.
Hoft, R. G.↗
Design Description of the X-33 Avionics Architecture
In this paper, we provide a design description of the X-33 avionics architecture. The X-33 is an autonomous Single Stage to Orbit (SSTO) launch vehicle currently being developed by Lockheed Martin for NASA as a technology demonstrator for the VentureStar Reusable Launch Vehicle (RLV). The X-33 avionics provides autonomous control of die vehicle throughout takeoff, ascent, descent, approach, landing, rollout, and vehicle safing. During flight the avionics provides communication to the range through uplinked commands and downlinked telemetry. During pre-launch and post-safing activities, the avionics provides interfaces to ground support consoles that perform vehicle flight preparations and maintenance. The X-33 Avionics is a hybrid of centralized and distributed processing elements connected by three dual redundant Mil-Std 1553 data buses. These data buses are controlled by a central processing suite located in the avionics bay and composed of triplex redundant Vehicle Mission Computers (VMCs). The VMCs integrate mission management, guidance, navigation, flight control, subsystem control and redundancy management functions. The vehicle sensors, effectors and subsystems are interfaced directly to the centralized VMCs as remote terminals or through dual redundant Data Interface Units (DIUs). The DIUs are located forward and aft of the avionics bay and provide signal conditioning, health monitoring, low level subsystem control and data interface functions. Each VMC is connected to all three redundant 1553 data buses for monitoring and provides a complete identical data set to the processing algorithms. This enables bus faults to be detected and reconfigured through a voted bus control configuration. Data is also shared between VMCs though a cross channel data link that is implemented in hardware and controlled by AlliedSignal's Fault Tolerant Executive (FTE). The FTE synchronizes processors within the VMC and synchronizes redundant VMCs to each other. The FTE provides an output-voting plane to detect, isolate and contain faults due to internal hardware or software faults and reconfigures the VMCs to accommodate these faults. Critical data in the 1553 messages are scheduled and synchronized to specific processing frames in order to minimize data latency. In order to achieve an open architecture, military and commercial off-the-shelf equipment is incorporated using common processors, standard VME backplanes and chassis, the VxWorks operating system, and MartixX for automatic code generation. The use of off-the-shelf tools and equipment helps reduce development time and enables software reuse. The open architecture allows for technology insertion, while the distributed modular elements allow for expansion to increased redundancy levels to meet the higher reliability goals of future RLVs.
Reichenfeld, Curtis J.↗
Evolution of Orion Mission Design for Exploration Mission 1 and 2
The evolving mission design and concepts of NASA's next steps have shaped Orion into the spacecraft that it is today. Since the initial inception of Orion, through the Constellation Program, and now in the Exploration Mission frame-work with the Space Launch System (SLS), each mission design concept and program goal have left Orion with a set of capabilities that can be utilized in many different mission types. Exploration Missions 1 and 2 (EM-1 and EM-2) have now been at the forefront of the mission design focus for the last several years. During that time, different Design Reference Missions (DRMs) were built, analyzed, and modified to solve or mitigate enterprise level design trades to ensure a viable mission from launch to landing. The resulting DRMs for EM-1 and EM-2 were then expanded into multi-year trajectory scans to characterize vehicle performance and Earth-Moon geometry trends. This provides Orion's subsystems with stressing reference trajectories to help design their system. Now that Orion has progressed through the Preliminary and Critical Design Re-views (PDR and CDR) there is a general shift in the focus of mission design from aiding the vehicle design to providing mission specific products needed for pre-flight and real time operations. Some of the mission specific products need-ed include analysis of steering law performance, inputs into navigational accura-cy assessments, abort options at any point in the mission for each valid trajecto-ry in the launch window, recontact avoidance between the upper stage and Orion post nominal separation, etc.
Gutkowski, Jeffrey P.↗
Optical Property Requirements for Glasses, Ceramics and Plastics in Spacecraft Window Systems
This is a preliminary draft of a standard published by the National Aeronautics and Space Administration (NASA) Johnson Space Center (JSC) that is intended to provide uniform window optical design requirements in support of the development of human-rated spaceflight hardware. The material covered in this standard is based on data from extensive testing by the Advanced Sensing and Optical Measurement Branch at NASA Langley Research Center, and compiled into requirements format by the NASA JSC Structural Engineering Division. At the time of this initial document release, a broader technical community has not reviewed this standard. The technical content of this standard is primarily based on the Constellation Program Orion Crew Exploration Vehicle Window Optical Properties Requirements, CxP 72407, Baseline. Unlike other optical requirements documents available for human rated spacecraft, this document includes requirements that ensure functionality for windows that contain glass/ceramic and/or plastic window substrate materials. These requirements were derived by measuring the optical properties of fused silica and aluminosilicate glass window assemblies and ensuring that the performance of any window assembly that includes a plastic pane or panes will meet the performance level of the all-glass assemblies. The resulting requirements are based upon the performance and parameter metrology testing of a variety of materials, including glass, transparent ceramics, acrylics, and polycarbonates. In general, these requirements are minimum specifications for each optical parameter in order to achieve the function specified for each functional category, A through D. Because acrylic materials perform at a higher level than polycarbonates in the optics regime, and CxP/Orion is planning to use acrylic in the Orion spacecraft, these requirements are based heavily on metrology from that material. As a result, two of the current Category D requirements for plastics are cited in such a way that will result in the screening out of polycarbonates. It is acknowledged that many polycarbonates can perform the functions of Category D, such as piloting and imagery with lens with apertures up to 25mm, without performance issues. Therefore, this forward warns users that certain requirements, such as birefringence and wavefront, for Category D plastics need to be revised to allow those polycarbonates that perform adequately in Category D to be accepted, while at the same time, screen out those materials that do not perform up to par. At the time of document release, the requirements in question have been identified by a TBD beside the proposed requirement criteria (which is based upon acrylic performance). Vehicles that are designed with acrylic materials for windowpanes are encouraged to use the values presented in this document for all requirements, in order to ensure adequate optical performance.
Lynda R Estes↗
Aerodynamic behavior of the Viking entry vehicle - Ground test and flight results
An extensive series of tests of the Viking entry vehicle flying in pure CO2 was conducted in a ballistic range at Ames Research Center. The primary purpose of these tests was to calibrate the aerodynamic lift and drag characteristics in order to allow the density, pressure, and temperature profiles of the Martian atmosphere to be determined from onboard instrumentation carried on Viking. Both the Viking 1 and Viking 2 entry vehicles performed flawlessly during entry and descent, and the atmosphere structure was deduced to an altitude of about 120 km. A description is given of the ballistic range tests and of the aerodynamic behavior of the full scale entry vehicles during entry into the Martian atmosphere. Some comparisons between ground test and flight results are shown.
Kirk, D. B.↗