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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 271 records · Page 15
All sourcesNASA NTRSDOE OSTI
NASA NTRS2005

Parachute Testing for the NASA X-38 Crew Return Vehicle

NASA's X-38 program was an in-house technology demonstration program to develop a Crew Return Vehicle (CRV) for the International Space Station capable of returning seven crewmembers to Earth when the Space Shuttle was not present at the station. The program, managed out of NASA's Johnson Space Center, was started in 1995 and was cancelled in 2003. Eight flights with a prototype atmospheric vehicle were successfully flown at Edwards Air Force Base, demonstrating the feasibility of a parachute landing system for spacecraft. The intensive testing conducted by the program included testing of large ram-air parafoils. The flight test techniques, instrumentation, and simulation models developed during the parachute test program culminated in the successful demonstration of a guided parafoil system to land a 25,000 Ib spacecraft. The test program utilized parafoils of sizes ranging from 750 to 7500 p. The guidance, navigation, and control system (GN&C) consisted of winches, laser or radar altimeter, global positioning system (GPS), magnetic compass, barometric altimeter, flight computer, and modems for uplink commands and downlink data. The winches were used to steer the parafoil and to perform the dynamic flare maneuver for a soft landing. The laser or radar altimeter was used to initiate the flare. In the event of a GPS failure, the software navigated by dead reckoning using the compass and barometric altimeter data. The GN&C test beds included platforms dropped from cargo aircraft, atmospheric vehicles released from a 8-52, and a Buckeye powered parachute. This paper will describe the test program and significant results.

Stein, Jenny M.↗
NASA NTRS2007

The Dangers of Failure Masking in Fault-Tolerant Software: Aspects of a Recent In-Flight Upset Event

On 1 August 2005, a Boeing Company 777-200 aircraft, operating on an international passenger flight from Australia to Malaysia, was involved in a significant upset event while flying on autopilot. The Australian Transport Safety Bureau's investigation into the event discovered that an anomaly existed in the component software hierarchy that allowed inputs from a known faulty accelerometer to be processed by the air data inertial reference unit (ADIRU) and used by the primary flight computer, autopilot and other aircraft systems. This anomaly had existed in original ADIRU software, and had not been detected in the testing and certification process for the unit. This paper describes the software aspects of the incident in detail, and suggests possible implications concerning complex, safety-critical, fault-tolerant software.

Johnson, C. W.↗
NASA NTRS2007

Integrated inertial stellar attitude sensor

An integrated inertial stellar attitude sensor for an aerospace vehicle includes a star camera system, a gyroscope system, a controller system for synchronously integrating an output of said star camera system and an output of said gyroscope system into a stream of data, and a flight computer responsive to said stream of data for determining from the star camera system output and the gyroscope system output the attitude of the aerospace vehicle.

Brady, Tye M.↗
NASA NTRS2008

Heating Augmentation for Short Hypersonic Protuberances

Computational aeroheating analyses of the Space Shuttle Orbiter plug repair models are validated against data collected in the Calspan University of Buffalo Research Center (CUBRC) 48 inch shock tunnel. The comparison shows that the average difference between computed heat transfer results and the data is about 9.5%. Using CFD and Wind Tunnel (WT) data, an empirical correlation for estimating heating augmentation on short hypersonic protuberances (k/delta less than 0.3) is proposed. This proposed correlation is compared with several computed flight simulation cases and good agreement is achieved. Accordingly, this correlation is proposed for further investigation on other short hypersonic protuberances for estimating heating augmentation.

Mazaheri, Ali R.↗
NASA NTRS2008

J-2X Abort System Development

The J-2X is an expendable liquid hydrogen (LH2)/liquid oxygen (LOX) gas generator cycle rocket engine that is currently being designed as the primary upper stage propulsion element for the new NASA Ares vehicle family. The J-2X engine will contain abort logic that functions as an integral component of the Ares vehicle abort system. This system is responsible for detecting and responding to conditions indicative of impending Loss of Mission (LOM), Loss of Vehicle (LOV), and/or catastrophic Loss of Crew (LOC) failure events. As an earth orbit ascent phase engine, the J-2X is a high power density propulsion element with non-negligible risk of fast propagation rate failures that can quickly lead to LOM, LOV, and/or LOC events. Aggressive reliability requirements for manned Ares missions and the risk of fast propagating J-2X failures dictate the need for on-engine abort condition monitoring and autonomous response capability as well as traditional abort agents such as the vehicle computer, flight crew, and ground control not located on the engine. This paper describes the baseline J-2X abort subsystem concept of operations, as well as the development process for this subsystem. A strategy that leverages heritage system experience and responds to an evolving engine design as well as J-2X specific test data to support abort system development is described. The utilization of performance and failure simulation models to support abort system sensor selection, failure detectability and discrimination studies, decision threshold definition, and abort system performance verification and validation is outlined. The basis for abort false positive and false negative performance constraints is described. Development challenges associated with information shortfalls in the design cycle, abort condition coverage and response assessment, engine-vehicle interface definition, and abort system performance verification and validation are also discussed.

Santi, Louis M.↗
NASA NTRS2008

A Stable Clock Error Model Using Coupled First and Second Order Gauss-Markov Processes

Long data outages may occur in applications of global navigation satellite system technology to orbit determination for missions that spend significant fractions of their orbits above the navigation satellite constellation(s). Current clock error models based on the random walk idealization may not be suitable in these circumstances, since the covariance of the clock errors may become large enough to overflow flight computer arithmetic. A model that is stable, but which approximates the existing models over short time horizons is desirable. A coupled first- and second-order Gauss-Markov process is such a model.

Carpenter, Russell↗
NASA NTRS2008

Heating Augmentation for Short Hypersonic Protuberances

Computational aeroheating analyses of the Space Shuttle Orbiter plug repair models are validated against data collected in the Calspan University of Buffalo Research Center (CUBRC) 48 inch shock tunnel. The comparison shows that the average difference between computed heat transfer results and the data is about 9:5%. Using CFD and Wind Tunnel (WT) data, an empirical correlation for estimating heating augmentation on short hyper- sonic protuberances (k/delta < 0.33) is proposed. This proposed correlation is compared with several computed flight simulation cases and good agreement is achieved. Accordingly, this correlation is proposed for further investigation on other short hypersonic protuberances for estimating heating augmentation.

Mazaheri, Alireza R.↗
NASA NTRS2007

Enhanced Master Controller Unit Tester

The Enhanced Master Controller Unit Tester (EMUT) software is a tool for development and testing of software for a master controller (MC) flight computer. The primary function of the EMUT software is to simulate interfaces between the MC computer and external analog and digital circuitry (including other computers) in a rack of equipment to be used in scientific experiments. The simulations span the range of nominal, off-nominal, and erroneous operational conditions, enabling the testing of MC software before all the equipment becomes available.

Benson, Patricia↗
NASA NTRS2009

Ares I Avionics Introduction

The Ares I is the next generation human-rated launcher for the United States Constellation program. This system is required to provide single fault tolerance within defined crew safety and mission reliability limits. As part of the effort to achieve those safety goals, Ares I includes an avionics subsystem built as a multistring, voting architecture. The avionics design draws upon experience gained from building fly-by-wire systems for Shuttle, X- 38, and Seawolf. Architectural drivers for the avionics design include using proven technologies with existing suppliers of space rated parts for critical functions (to reduce overall development risk), easing the software development effort by using an off-theshelf, DO-178B certifiable, ARINC-653 operating system in the main flight computers, minimizing mutual data and power connections that might lead to a common-mode hardware failure of the redundant avionics strings, and centralizing overall Ares I command & control within the Upper Stage.

Marchant, Christopher C.↗
NASA NTRS2011

FPGA for Power Control of MSL Avionics

A PLGT FPGA (Field Programmable Gate Array) is included in the LCC (Load Control Card), GID (Guidance Interface & Drivers), TMC (Telemetry Multiplexer Card), and PFC (Pyro Firing Card) boards of the Mars Science Laboratory (MSL) spacecraft. (PLGT stands for PFC, LCC, GID, and TMC.) It provides the interface between the backside bus and the power drivers on these boards. The LCC drives power switches to switch power loads, and also relays. The GID drives the thrusters and latch valves, as well as having the star-tracker and Sun-sensor interface. The PFC drives pyros, and the TMC receives digital and analog telemetry. The FPGA is implemented both in Xilinx (Spartan 3- 400) and in Actel (RTSX72SU, ASX72S). The Xilinx Spartan 3 part is used for the breadboard, the Actel ASX part is used for the EM (Engineer Module), and the pin-compatible, radiation-hardened RTSX part is used for final EM and flight. The MSL spacecraft uses a FC (Flight Computer) to control power loads, relays, thrusters, latch valves, Sun-sensor, and star-tracker, and to read telemetry such as temperature. Commands are sent over a 1553 bus to the MREU (Multi-Mission System Architecture Platform Remote Engineering Unit). The MREU resends over a remote serial command bus c-bus to the LCC, GID TMC, and PFC. The MREU also sends out telemetry addresses via a remote serial telemetry address bus to the LCC, GID, TMC, and PFC, and the status is returned over the remote serial telemetry data bus.

Wang, Duo↗
NASA NTRS2011

Kalman Orbit Optimized Loop Tracking

Under certain conditions of low signal power and/or high noise, there is insufficient signal to noise ratio (SNR) to close tracking loops with individual signals on orbiting Global Navigation Satellite System (GNSS) receivers. In addition, the processing power available from flight computers is not great enough to implement a conventional ultra-tight coupling tracking loop. This work provides a method to track GNSS signals at very low SNR without the penalty of requiring very high processor throughput to calculate the loop parameters. The Kalman Orbit-Optimized Loop (KOOL) tracking approach constitutes a filter with a dynamic model and using the aggregate of information from all tracked GNSS signals to close the tracking loop for each signal. For applications where there is not a good dynamic model, such as very low orbits where atmospheric drag models may not be adequate to achieve the required accuracy, aiding from an IMU (inertial measurement unit) or other sensor will be added. The KOOL approach is based on research JPL has done to allow signal recovery from weak and scintillating signals observed during the use of GPS signals for limb sounding of the Earth s atmosphere. That approach uses the onboard PVT (position, velocity, time) solution to generate predictions for the range, range rate, and acceleration of the low-SNR signal. The low- SNR signal data are captured by a directed open loop. KOOL builds on the previous open loop tracking by including feedback and observable generation from the weak-signal channels so that the MSR receiver will continue to track and provide PVT, range, and Doppler data, even when all channels have low SNR.

Young, Lawrence E.↗
NASA NTRS2012

On Component Reliability and System Reliability for Space Missions

This paper is to address the basics, the limitations and the relationship between component reliability and system reliability through a study of flight computing architectures and related avionics components for NASA future missions. Component reliability analysis and system reliability analysis need to be evaluated at the same time, and the limitations of each analysis and the relationship between the two analyses need to be understood.

Chen, Yuan↗
NASA NTRS2012

Pose Measurement Performance of the Argon Relative Navigation Sensor Suite in Simulated Flight Conditions

Argon is a flight-ready sensor suite with two visual cameras, a flash LIDAR, an on- board flight computer, and associated electronics. Argon was designed to provide sensing capabilities for relative navigation during proximity, rendezvous, and docking operations between spacecraft. A rigorous ground test campaign assessed the performance capability of the Argon navigation suite to measure the relative pose of high-fidelity satellite mock-ups during a variety of simulated rendezvous and proximity maneuvers facilitated by robot manipulators in a variety of lighting conditions representative of the orbital environment. A brief description of the Argon suite and test setup are given as well as an analysis of the performance of the system in simulated proximity and rendezvous operations.

Galante, Joseph M.↗
NASA NTRS2011

Mars Exploration Rover Heatshield Observation Campaign

For the first time ever, engineers were able to observe a heatshield on the surface of another planet after a successful entry through the atmosphere. A three-week heatshield observation campaign was conducted in December 2004 after the Mars Exploration Rover Opportunity exited "Endurance Crater." By utilizing the rover's scientific instruments, data was collected to make a qualitative assessment of the performance of the heatshield. This data was gathered to gain a better understanding of how the heatshield performed during entry through the Martian atmosphere. In addition, this unprecedented look at the heatshield offered engineers the opportunity to assess if any unexpected anomalies occurred. Once a survey of the heatshield debris was completed, multiple targets of interest were chosen for the collection of imaging data. This data was then used to assess the char depth of the thermal protection material, which compared well with design and post-flight computational predictions. Extensive imaging data was collected and showed the main seal in pristine conditions, and no observable indications of structure overheating. Additionally, unexpected vehicle dynamics during the atmospheric entry were explained by the observation of thermal blanket remnants attached to the heatshield.

Thermal Protection System (TPS)↗
NASA NTRS2013

Trajectory Reconstruction and Uncertainty Analysis Using Mars Science Laboratory Pre-Flight Scale Model Aeroballistic Testing

As part of the Mars Science Laboratory (MSL) trajectory reconstruction effort at NASA Langley Research Center, free-flight aeroballistic experiments of instrumented MSL scale models was conducted at Aberdeen Proving Ground in Maryland. The models carried an inertial measurement unit (IMU) and a flush air data system (FADS) similar to the MSL Entry Atmospheric Data System (MEADS) that provided data types similar to those from the MSL entry. Multiple sources of redundant data were available, including tracking radar and on-board magnetometers. These experimental data enabled the testing and validation of the various tools and methodologies that will be used for MSL trajectory reconstruction. The aerodynamic parameters Mach number, angle of attack, and sideslip angle were estimated using minimum variance with a priori to combine the pressure data and pre-flight computational fluid dynamics (CFD) data. Both linear and non-linear pressure model terms were also estimated for each pressure transducer as a measure of the errors introduced by CFD and transducer calibration. Parameter uncertainties were estimated using a "consider parameters" approach.

Lugo, Rafael A.↗
NASA NTRS2013

Discrete Data Qualification System and Method Comprising Noise Series Fault Detection

A Sensor Data Qualification (SDQ) function has been developed that allows the onboard flight computers on NASA s launch vehicles to determine the validity of sensor data to ensure that critical safety and operational decisions are not based on faulty sensor data. This SDQ function includes a novel noise series fault detection algorithm for qualification of the output data from LO2 and LH2 low-level liquid sensors. These sensors are positioned in a launch vehicle s propellant tanks in order to detect propellant depletion during a rocket engine s boost operating phase. This detection capability can prevent the catastrophic situation where the engine operates without propellant. The output from each LO2 and LH2 low-level liquid sensor is a discrete valued signal that is expected to be in either of two states, depending on whether the sensor is immersed (wet) or exposed (dry). Conventional methods for sensor data qualification, such as threshold limit checking, are not effective for this type of signal due to its discrete binary-state nature. To address this data qualification challenge, a noise computation and evaluation method, also known as a noise fault detector, was developed to detect unreasonable statistical characteristics in the discrete data stream. The method operates on a time series of discrete data observations over a moving window of data points and performs a continuous examination of the resulting observation stream to identify the presence of anomalous characteristics. If the method determines the existence of anomalous results, the data from the sensor is disqualified for use by other monitoring or control functions.

Fulton, Christopher↗
NASA NTRS2013

Project Morpheus: Morpheus 1.5A Lander Failure Investigation Results

On August 9, 2012 the Morpheus 1.5A vehicle crashed shortly after lift off from the Kennedy Space Center. The loss was limited to the vehicle itself which was pre-declared to be a test failure and not a mishap. The Morpheus project is demonstrating advanced technologies for in space and planetary surface vehicles including: autonomous flight control, landing site hazard identification and safe site selection, relative surface and hazard navigation, precision landing, modular reusable flight software, and high performance, non-toxic, cryogenic liquid Oxygen and liquid Methane integrated main engine and attitude control propulsion system. A comprehensive failure investigation isolated the fault to the Inertial Measurement Unit (IMU) data path to the flight computer. Several improvements have been identified and implemented for the 1.5B and 1.5C vehicles.

Devolites, Jennifer L.↗
NASA NTRS2014

Lessons Learned in the First Year Operating Software Defined Radios in Space

Operating three unique software defined radios (SDRs) in a space environment aboard the Space Communications and Navigation (SCaN) Testbed for over one year has provided an opportunity to gather knowledge useful for future missions considering using software defined radios. This paper provides recommendations for the development and use of SDRs, and it considers the details of each SDRs approach to software upgrades and operation. After one year, the SCaN Testbed SDRs have operated for over 1000 hours. During this time, the waveforms launched with the SDR were tested on-orbit to assure that they operated in space at the same performance level as on the ground prior to launch to obtain an initial on-orbit performance baseline. A new waveform for each SDR has been developed, implemented, uploaded to the flight system, and tested in the flight environment. Recommendations for SDR-based missions have been gathered from early development through operations. These recommendations will aid future missions to reduce the cost, schedule, and risk of operating SDRs in a space environment. This paper considers the lessons learned as they apply to SDR pre-launch checkout, purchasing space-rated hardware, flexibility in command and telemetry methods, on-orbit diagnostics, use of engineering models to aid future development, and third-party software. Each SDR implements the SCaN Testbed flight computer command and telemetry interface uniquely, allowing comparisons to be drawn. The paper discusses the lessons learned from these three unique implementations, with suggestions on the preferred approach. Also, results are presented showing that it is important to have full system performance knowledge prior to launch to establish better performance baselines in space, requiring additional test applications to be developed pre-launch. Finally, the paper presents the issues encountered with the operation and implementation of new waveforms on each SDR and proposes recommendations to avoid these issues.

radio equipment↗