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.
Bench-scale and pilot plant development of solvent refined coal (SRC) processes is reviewed. Large SRC demonstration plants are described. Commercialization of the process is envisioned for the 1980's.
We report Swift Burst Alert Telescope (BAT) observations of the X-ray flash (XRF) XRF 050416A. The fluence ratio between the 15-25 and 25-50 keV energy bands of this event is 1.5, thus making it the softest gamma-ray burst (GRB) observed by BAT so far. The spectrum is well fitted by a Band function with E(sup obs)(sub peak) of 15.0(sup +2.3)(sub -2.7) keV. Assuming the redshift of the host galaxy (z = 0.6535), the isotropic equivalent radiated energy E(sub iso) and the peak energy at the GRB rest frame (E(sup src)(sub peak)) of XRF 050416A are not only consistent with the correlation found by Amati et al. and extended to XRFs by Sakamoto et al. but also fill in the gap of this relation around the 30-80 keV range of E(sup src)(sub peak). This result tightens the validity of the E(sup src)(sub Peak)-E(sup src)(sub peak) relation from XRFs to GRBs. We also find that the jet break time estimated using the empirical relation between E(sup src)(sub peak) and the collimation corrected energy E(sub gamma), is inconsistent with the afterglow observation by the Swift X-Ray Telescope. This could be due to the extra external shock emission overlaid around the jet break time or to the nonexistence of a jet break feature for XRFs, which might be a further challenge for GRB jet emission models and XRF/GRB unification scenarios.
Collectors at 1173K (900 C) were exposed to the combustion products of a Mach 0.3 burner rig fueled with various industrial turbine liquid fuels from solvent refined coals. Four fuels were employed: a naphtha, a light oil, a wash solvent and a mid-heavy distillate blend. The response of four superalloys (IN-100, U 700, IN 792 and M-509) to exposure to the combustion gases from the SRC-2 naphtha and resultant deposits was also determined. The SRC-2 fuel analysis and insights obtained during the combustion experience are discussed. Particular problems encountered were fuel instability and reactions of the fuel with hardware components. The major metallic elements which contributed to the deposits were copper, iron, chromium, calcium, aluminum, nickel, silicon, titanium, zinc, and sodium. The deposits were found to be mainly metal oxides. An equilibrium thermodynamic analysis was employed to predict the chemical composition of the deposits. The agreement between the predicted and observed compounds was excellent. No hot corrosion was observed. This was expected because the deposits contained very little sodium or potassium and consisted mainly of the unreactive oxides. However, the amounts of deposits formed indicated that fouling is a potential problem with the use of these fuels.
The Saskatchewan Research Council (SRC) collected surface meteorological and radiation data from December 1993 until December 1996. The data set comprises Suite A (meteorological and energy balance measurements) and Suite B (diffuse solar and longwave measurements) components. Suite A measurements were taken at each of ten sites, and Suite B measurements were made at five of the Suite A sites. The data cover an approximate area of 500 km (North-South) by 1000 km (East-West) (a large portion of northern Manitoba and northern Saskatchewan). The measurement network was designed to provide researchers with a sufficient record of near-surface meteorological and radiation measurements. The data are provided in tabular ASCII files, and were collected by Aircraft Flux and Meteorology (AFM)-7. The surface meteorological and radiation data are available from the Earth Observing System Data and Information System (EOSDIS) Oak Ridge National Laboratory (ORNL) Distributed Active Archive Center (DAAC). The data files are available on a CD-ROM (see document number 20010000884).
The primary mission of the NASA Mars Sample Return (MSR) Campaign is to return samples of the Martian surface to Earth for scientific study. As part of this campaign, NASA is developing a Mars Ascent Vehicle (MAV). This vehicle must survive an approximate two year journey to the Martian surface as a payload aboard a separate lander spacecraft. After residing on the surface for another year, the MAV will carry a payload of samples into orbit. From there, following ejection from the MAV, the samples will rendezvous with an Earth return spacecraft for capture, and ultimately, return to Earth.The design of the MAV represents a number of unique challenges, as no launch vehicle has ever left the surface of a planet other than Earth. Although conceptual designs for a MAV have been in various levels of development since the 1970s, none have achieved the level of fidelity and support that exists in the current MSR-MAV design. Early MSR-MAV concept studies examined multiple methods of propulsion, ultimately deciding that a Two Stage to Orbit (TSTO) solid propulsion vehicle would provide the most capable performance in a Martian environment. Following this key architecture decision, the vehicle design was further matured to a Solid-Solid Guided-Guided (SSGG) architecture for NASA Key Decision Point A (KDP-A). Although the SSGG design was able to meet all mission constraints, concerns were raised regarding limited mass margin on other elements of the MSR campaign at such an early phase. A design challenge was issued to reduce MAV total mass by as much as possible. It was ultimately determined that by moving a number of components of the vehicle second stage to the first stage, the overall vehicle mass could be reduced significantly. The new design featured a much smaller and completely unguided second stage. This paper describes the resultant Solid-Solid Guided-Unguided (SSGU) MAV design concept developed as part of the Systems Requirement Cycle (SRC). This design was developed primarily by NASA Marshall Space Flight Center (MSFC), in association with NASA Jet Propulsion Laboratory (JPL) and NASA Langley Research Center (LaRC). The TSTO vehicle includes one solid rocket motor per stage. As the vehicle second stage is unguided, it features spin-stabilization to maintain vehicle stability during flight. An electromechanically actuated Thrust Vector Control (TVC) and a monopropellant Reaction Control System (RCS) is employed for active guidance on the first stage. The vehicle is designed to deliver up to 0.47kg of Martian samples to a Mars circular orbit of 380km at 27° inclination. Due to the extremely unique design constraints of this mission, and a recent transition to a Risk Class A posture, the MAV team was compelled to devise unconventional solutions to the vehicle design. The detailed design and analysis of these subsystems and the vehicle as a whole are discussed in this paper relative to all of the engineering disciplines involved.
The primary mission of the NASA Mars Sample Return (MSR) Campaign is to return samples of the Martian surface to Earth for scientific study. As part of this campaign, NASA is developing a Mars Ascent Vehicle (MAV). This vehicle must survive an approximate two-year journey to the Martian surface as a payload aboard a separate lander spacecraft. After residing on the surface for another year, the MAV will carry a payload of samples into orbit. From there, following ejection from the MAV, the samples will rendezvous with an Earth return spacecraft for capture, and ultimately, return to Earth. The design of the MAV represents a number of unique challenges, as no launch vehicle has ever left the surface of a planet other than Earth. Although conceptual designs for a MAV have been in various levels of development since the 1970s, none have achieved the level of fidelity and support that exists in the current MSR-MAV design. Early MSR-MAV concept studies examined multiple methods of propulsion, ultimately deciding that a Two-Stage-to-Orbit (TSTO) solid propulsion vehicle would provide the most capable performance in a Martian environment. Following this key architecture decision, the vehicle design was further matured to a Solid-Solid Guided-Guided (SSGG) architecture for NASA Key Decision Point A (KDP-A). Although the SSGG design was able to meet all mission constraints, concerns were raised regarding limited mass margin on other elements of the MSR campaign at such an early phase. A design challenge was issued to reduce MAV total mass by as much as possible. It was ultimately determined that by moving a number of components of the vehicle second stage to the first stage, the overall vehicle mass could be reduced significantly. The new design features a much smaller and completely unguided second stage. This paper describes the resultant Solid-Solid Guided-Unguided (SSGU) MAV design concept developed as part of the Systems Requirement Cycle (SRC). This design was developed primarily by NASA Marshall Space Flight Center (MSFC), in association with NASA Jet Propulsion Laboratory (JPL) and NASA Langley Research Center (LaRC). The TSTO vehicle includes one solid rocket motor per stage. As the vehicle second stage is unguided, it features spin-stabilization to maintain vehicle stability during flight. An electromechanically actuated Thrust Vector Control (TVC) system and a monopropellant Reaction Control System (RCS) are employed for active guidance on the first stage. The vehicle is designed to deliver up to 0.47kg of Martian samples to a Mars circular orbit of 380km at 27° inclination. Due to the extremely unique design constraints of this mission, and a recent transition to a Risk Class A posture, the MAV team was compelled to devise unconventional solutions to the vehicle design. The detailed design and analysis of these subsystems and the vehicle as a whole are discussed in this paper relative to all of the engineering disciplines involved.
The objective of the NASA Mars Sample Return (MSR) Campaign is to collect samples from the surface of Mars and return them to Earth for scientific research. The Mars Ascent Vehicle (MAV) will be integrated into a larger Mars Sample Retrieval Lander (SRL) for transit to and storage on Mars. After all Martian samples have been collected and loaded into the MAV payload assembly, MAV will deliver the samples from the Martian surface to Mars orbit. A separate spacecraft, the Earth Return Orbiter (ERO) will retrieve the samples from Mars orbit and return them to Earth. To address common systems engineering challenges associated with using traditional systems engineering practices on complex projects, the MAV systems engineering team has explored implementation of Model-Based Systems Engineering (MBSE) tools and languages. This paper describes the current state of implementation and development of the MAV MBSE model with the Systems Modeling Language (SysML) within the scope of the MAV Systems Requirement Cycle (SRC) systems engineering workflow. The MAV MBSE model has been developed within Magic Draw – a SysML editor commonly used to implement MBSE. The MAV MBSE model has been used to develop mission phase functional flow diagrams for the Concept of Operations, decompose mission to vehicle subsystem functions, develop a functional decomposition, derive functional requirements, trace requirements up to customer-imposed requirements, trace requirements within MAV requirement space, identify requirements trace gaps, define and map the physical design space architecture, allocate requirements to subsystems, develop validation items, define assembly, integration, and test (AI&T) operations, and trace these items across driving goals to develop an integrated digital thread of systems engineering information used to drive design specifications, decision making, and ultimately design verification and validation. Findings and results associated with implementing MBSE in these ways, alongside traditional methods will be discussed.
The objective of the NASA Mars Sample Return (MSR) Campaign is to collect samples from the surface of Mars and return them to Earth for scientific research. The Mars Ascent Vehicle (MAV) will be integrated into a larger Mars Sample Retrieval Lander (SRL) for transit to and storage on Mars. After all Martian samples have been collected and loaded into the MAV payload assembly, MAV will deliver the samples from the Martian surface to Mars orbit. A separate spacecraft, the Earth Return Orbiter (ERO) will retrieve the samples from Mars orbit and return them to Earth. To address common systems engineering challenges associated with using traditional systems engineering practices on complex projects, the MAV systems engineering team has explored implementation of Model-Based Systems Engineering (MBSE) tools and languages. This paper describes the current state of implementation and development of the MAV MBSE model with the Systems Modeling Language (SysML) within the scope of the MAV Systems Requirement Cycle (SRC) systems engineering workflow. The MAV MBSE model has been developed within Magic Draw – a SysML editor commonly used to implement MBSE. The MAV MBSE model has been used to develop mission phase functional flow diagrams for the Concept of Operations, decompose mission to vehicle subsystem functions, develop a functional decomposition, derive functional requirements, trace requirements up to customer-imposed requirements, trace requirements within MAV requirement space, identify requirements trace gaps, define and map the physical design space architecture, allocate requirements to subsystems, develop validation items, define assembly, integration, and test (AI&T) operations, and trace these items across driving goals to develop an integrated digital thread of systems engineering information used to drive design specifications, decision making, and ultimately design verification and validation. Findings and results associated with implementing MBSE in these ways, alongside traditional methods will be discussed.
Small venthole drilled in semisealed silicon-controlled rectifier (SCR) cavity eliminates entrapped helium. Although these devices show slightly greater leak than those before lead installation, it is now possible to distinguish device with good hermetic seal from defective one.
Phenolic Impregnated Carbon Ablator (PICA) was developed at NASA Ames Research Center. As a thermal protection material, PICA has the advantages of being able to withstand high heat fluxes with a relatively low density. This ablative material was used as the forebody heat shield material for the Stardust sample return capsule, which re-entered the Earths atmosphere in 2006. Based on PICA, SpaceX developed a variant, PICA-X, and used it as the heat shield material for its Dragon spacecraft, which successfully orbited the Earth and re-entered the atmosphere during the COTS Demo Flight 1 in 2010. Post-flight analysis was previously performed on the Stardust PICA heat shield material. Similarly, a near-stagnation core was obtained from the post-flight Dragon 1 heat shield, which was retrieved from the Pacific Ocean. Materials testing and analyses were performed on the core to evaluate its ablation performance and post-flight properties. Comparisons between PICA and PICA-X are made where applicable. Stardust and Dragon offer rare opportunities to evaluate materials post-flight - this data is beneficial in understanding material performance and also improves modeling capabilities.
Explore the source record for details and available documents.
Extraterrestrial sample return is a growing component of solar system exploration. Currently, four missions, Stardust, 1 Muses-C, 2 Genesis, and Mars Sample Return, are under development that employ sample return as a prime component of the mission architecture. Respectively, these missions will return samples from the tail of a comet, an asteroid, the solar wind, and, Mars. An important component of these missions and the focus of this paper is the design of the sample return capsule (SRC). The purpose of the SRC is to safely return to Earth any gathered samples for terrestrial analysis. The two major design constraints for any SRC are as follows: 1) it must be able to survive a high-speed Earth entry (11 km/s to as a high as 15 km/s), 2) the mass of the SRC must be as small as possible. Because the SRC mass is carried from Earth to the sample sight and back, the SRC mass is a strong driver in the mission mass budget. Further, for the Mars Sample Return Capsule, planetary protection is another constraint. For this capsule, the probability of planetary contamination at Earth due to an SRC failure at entry must be minimal. For an SRC, a possible failure mechanism is severe local heating as a result of cavities and or protuberances in the SRC forebody heatshield. For example, the Apollo Command Module had a number of cavities and protuberances as part of the baseline designs Wind-tunnel tests of models containing small cavities and protuberances showed severe local heating augmentations in the vicinity of these surface discontinuities.4-5 As another example, the Genesis SRC forebody heat-shield contains penetrations (cavities) to mount the vehicle to the carrier bus. It is expected that these penetrations will also experience a severe local heating environment. A concern is that the large thermal gradients may produce sufficient thermal stress to cause local mechanical failure of the heatshield. Penetrations to the forebody heat-shield can also result from damage at vehicle integration, during launch, or during transportation of the sample return capsule from earth to the sample site and back. For example, the Starting SRC was damaged near the shoulder during the heatshield integration process producing a local surface discontinuity. Also, the Starting SRC traverses through the tail of a comet and is in space for 7 years. Thus, damage to the heatshield as a result of micrometeroid impact is a concern. Finally, it is difficult to characterize the effects of these potential heatshield singularities with ground-test facilities. Either detailed simulation or a dedicated flight test is required.
The Japan Aerospace Exploration Agency (JAXA) recently completed their Hayabusa asteroid exploration mission. Launched in 2003, Hayabusa made contact with, and retrieved a sample from, the near-Earth asteroid Itokawa in 2005. The sample return capsule (SRC) re-entered over the Woomera Test Range (WTR) in southern Australia on June 13, 2010, at approximately 11:21 pm local time (09:51 UTC). The SRC re-entry velocity was 12.2 km/s, making it the second-fastest Earth return velocity behind NASA s Stardust sample return capsule re-entry in 2006. From a space technology development perspective, Hayabusa s re-entry functioned as a rare flight experiment of an entry vehicle and its thermal protection system. In collaboration with the SETI Institute, NASA deployed its DC-8 airborne laboratory and a team of international researchers to Australia to observe the re-entry of the SRC. The use of an airborne platform enables observation above most clouds and weather and greatly diminishes atmospheric absorption of the optical signals. The DC-8 s flight path was engineered and flown to provide a view of the spacecraft that bracketed the heat pulse to the capsule. A suite of imaging instruments on board the DC-8 successfully recorded the luminous portion of the re-entry event. For approximately 70 seconds, the spectroscopic and radiometric instruments acquired images and spectra of the capsule, its wake, and destructive re-entry of the spacecraft bus. Figure 1 shows a perspective view of the WTR, the SRC re-entry trajectory, and the flight path of the DC-8. The SRC was jettisoned from the spacecraft bus approximately 3 hours prior to entry interface. Due to thruster failures on the spacecraft, it could not be diverted from the entry path and followed the trajectory of the SRC, where it burned up in the atmosphere between approximately 100 and 50 km altitude. Fortuitously, the separation distance between the spacecraft and SRC was sufficient to clearly resolve the SRC from the debris field of the burning spacecraft. Figure 2 shows a frame from a high-definition television camera on board the aircraft and denotes the locations of the SRC and spacecraft bus debris.
The Stardust Sample Return Capsule separated from the main spacecraft four hours prior to atmospheric entry. Between this time and the time at which the SRC touched down at the Utah Test and Training Range, two organizations external to JPL were involved in tracking the Sample Return Capsule. Orbit determination for the Stardust spacecraft during deep space cruise, the encounters of asteroid Annefrank and comet Wild 2, and the final approach to Earth used X-band radio metric Doppler and range data obtained through the Deep Space Network. The SRC lacked the electronics needed for coherently transponded radio metric tracking, so the DSN was not able to track the SRC after it separated from the main spacecraft. Although the expected delivery accuracy at atmospheric entry was well within the capability needed to target the SRC to the desired ground location, it was still desirable to obtain direct knowledge of the SRC trajectory in case of anomalies. For this reason U.S. Strategic Command was engaged to track the SRC between separation and atmospheric entry. Once the SRC entered the atmosphere, ground sensors at UTTR were tasked to acquire the descending SRC and maintain track during the descent in order to determine the landing location, to which the ground recovery team was then directed. This paper discusses organizational interfaces, data products, and delivery schedules, and the actual tracking operations are described.
On January 7, 2001, the Genesis spacecraft lifted off from Cape Canaveral. Its mission was to collect solar wind samples and return those samples to Earth for detailed analysis by scientists. The mission proceeded successfully for three-and-a-half years. On September 8, 2004, the spacecraft approached Earth, pointed the Sample Return Capsule (SRC) at its entry target, and then fired pyros that jettisoned the SRC. The SRC carried the valuable samples collected over the prior 29 months. The SRC also contained the requisite hardware (mechanisms, parachutes, and electronics) to manage the process of entry, descent, and landing (EDL). After entering Earth’s atmosphere, the SRC was expected to open a drogue parachute. This should have been followed by a pyro event to release the drogue chute, and then by a pyro event to deploy the main parachute at an approximate elevation of 6.7 kilometers. As the SRC descended to the Utah landing site, helicopters were in position to capture the SRC before the capsule touched down. On September 8, 2004, observers of the SRC’s triumphant return became concerned as the NASA announcer fell silent, and then became even more alarmed as they watched the spacecraft tumble as it streaked across the sky. Long-distance cameras clearly showed that the drogue parachute had not deployed properly.
On January 7, 2001, the Genesis spacecraft lifted off from Cape Canaveral. Its mission was to collect solar wind samples and return those samples to Earth for detailed analysis by scientists. The mission proceeded successfully for three-and-a-half years. On September 8, 2004, the spacecraft approached Earth, pointed the Sample Return Capsule (SRC) at its entry target, and then fired pyros that jettisoned the SRC. The SRC carried the valuable samples collected over the prior 29 months. The SRC also contained the requisite hardware (mechanisms, parachutes, and electronics) to manage the process of entry, descent, and landing (EDL). After entering Earthas atmosphere, the SRC was expected to open a drogue parachute. This should have been followed by a pyro event to release the drogue chute, and then by a pyro event to deploy the main parachute at an approximate elevation of 6.7 kilometers. As the SRC descended to the Utah landing site, helicopters were in position to capture the SRC before the capsule touched down. On September 8, 2004, observers of the SRCas triumphant return became concerned as the NASA announcer fell silent, and then became even more alarmed as they watched the spacecraft tumble as it streaked across the sky. Long-distance cameras clearly showed that the drogue parachute had not deployed properly. On September 9, 2004, General Eugene Tattini, Deputy Director of the Jet Propulsion Laboratory formed a Failure Review Board (FRB). This board was charged with investigating the cause of the Genesis mishap in close concert with the NASA Mishap Investigation Board (MIB). The JPL-FRB was populated with experts from within and external to the Jet Propulsion Laboratory. The JPL-FRB participated with the NASA-MIB through all phases of the investigation, working jointly and concurrently as one team to discover the facts of the mishap.
The purpose of this paper was to provide insight into the nature of response selection by reviewing the literature on stimulus-response compatibility (SRC) effects and the psychological refractory period (PRP) effect individually and jointly. The empirical findings and theoretical explanations of SRC effects that have been studied within a single-task context suggest that there are two response-selection routes-automatic activation and intentional translation. In contrast, all major PRP models reviewed in this paper have treated response selection as a single processing stage. In particular, the response-selection bottleneck (RSB) model assumes that the processing of Task 1 and Task 2 comprises two separate streams and that the PRP effect is due to a bottleneck located at response selection. Yet, considerable evidence from studies of SRC in the PRP paradigm shows that the processing of the two tasks is more interactive than is suggested by the RSB model and by most other models of the PRP effect. The major implication drawn from the studies of SRC effects in the PRP context is that response activation is a distinct process from final response selection. Response activation is based on both long-term and short-term task-defined S-R associations and occurs automatically and in parallel for the two tasks. The final response selection is an intentional act required even for highly compatible and practiced tasks and is restricted to processing one task at a time. Investigations of SRC effects and response-selection variables in dual-task contexts should be conducted more systematically because they provide significant insight into the nature of response-selection mechanisms.
The IMAG Pilot Study, recently completed at the University of Texas Medical Branch, filled in the second major gap in knowledge standing in the way of development of a practical Short Radius Centrifuge (SRC) and the use of Artificial Gravity (AG) as a multi-system countermeasure to combat the deconditioning associated with extended weightlessness. (The first challenge, to adapt rapidly rotating subjects to permit unlimited head movements without excessive motion sickness, was achieved in a series of studies at MIT involving incremental increases in head and centrifuge velocity.) It remained to be demonstrated that intermittent exposure to AG, at only one hour per day for 21 days, would have any positive effect on slowing or eliminating of deconditioning. Bed-rested normal subjects were used as a ground analog for astronauts in weightlessness. The results are clearly positive for the key physiological systems of interest: cardiovascular, muscle, and bone. No functionally relevant changes were observed in immune, cognitive, or sensory-motor function. Furthermore, we found that our initial concerns about the inability of deconditioned subjects to withstand daily centrifugation without syncope were misplaced. These encouraging initial results clearly support the further development of AG protocols. We recommend, as the next steps, the integration of a controlled exercise device on the SRC to determine the synergy between AG and exercise. Coupled with appropriate exercise device(s) the AG protocol will be tuned to-ward an optimal prescription for minimum exposure duration and frequency, maximum AG level and SRC speed.. Performance of these next steps will require extensive use of bed-rest/centrifuge facilities and eventually validation using an SRC in space. A space SRC could be placed in the ISS or on a planetary surface.