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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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NASA NTRS1998

Behavior and Performance Project

Behavior is an overt manifestation of underlying biology. As such, alterations in biological systems that result from spaceflight would be expected to evidence themselves in subtle or even pronounced changes in the behavior of that organism. These alterations in visible behavior may then indicate or even be diagnostic of alterations in the physical well-being of humans and other animals as they adapt to space environments or readapt to Earth--alterations that might not otherwise be detected without relatively invasive and frequently expensive procedures. Moreover, behavior and performance constitute the central standard for evaluating the mission success of spaceflight ventures. The success or failure of any excursion into space is primarily indexed by the ability of astronauts and cosmonauts to perform the tasks and experiments of any particular mission, to land a re-entry vehicle after extended exposure to microgravity (as with the space shuttle), or to make time-critical and life-saving repairs or other decisions while in orbit (as with the recent Mir events). Thus, understanding the effects spaceflight on behavior and performance is inherently important, in addition to those insights that behavior can provide about the physiological consequences of space adaptation.

Washburn, David A.↗
NASA NTRS1972

Space Tug Aerobraking Study. Volume 2: Technical

The feasibility and practicality of employing an aerobraking trajectory for return of the reusable Space Tug from geosynchronous and other high energy missions was investigated. The aerobraking return trajectory modes from high orbits employ transfer ellipses which have low perigee altitudes wherein the earth's sensible atmosphere provides drag to reduce the Tug descent delta velocity requirements and thus decrease the required return trip propulsive energy. An aerobraked Space Tug, sized to the Space Shuttle payload capability and dimensional constraints, can accomplish 95 percent of the geosynchronous missions with a single Shuttle/Tug launch per mission. Aerodynamics, aerothermodynamics, trajectory, quidance and control, configuration concepts, materials, weights and performance parameters were identified. Sensitivities to trajectory uncertainties, atmospheric anomalies and re-entry environments were determined. New technology requirements and future studies required to further enhance the aerobraking potential were identified.

Corso, C. J.↗
NASA NTRS2001

Compton Gamma Ray Observatory: Lessons Learned in Propulsion

The Compton Gamma Ray Observatory was the second of NASA's Great Observatories. At 17 1/2 tons. it was the heaviest astrophysical payload ever flown at the time of its launch on April 5, 1991 aboard the Space Shuttle. During initial, on-orbit priming of the spacecraft's monopropellant hydrazine propulsion system, a severe waterhammer transient was experienced. At that time, anomalous telemetry readings were received from on-board propulsion system instrumentation. This led to ground analyses and laboratory investigations as to the root cause of the waterhammer, potential damage to system integrity and functionality, and risks for switching from the primary (A-side) propulsion system to the redundant (B-side) system. The switchover to B-side was ultimately performed successfully and the spacecraft completed its basic and extended missions in this configuration. Nine years later, following a critical control gyroscope failure, Compton was safely deorbited and re-entered the Earth's atmosphere on June 4, 2000. Additional risk assessments concerning viability of A- and B-sides were necessary to provide confidence in attitude and delta-V authority and reliability to manage the precisely controlled reentry. This paper summarizes the design and operation of the propulsion system used on the spacecraft and provides "lessons learned" from the system engineering investigations into the propellant loading procedures, the initial priming anomaly, mission operations, and the commanded re-entry following the gyro failure.

Dressler, G. A.↗
NASA NTRS1994

Plasma Heating: An Advanced Technology

The Mercury and Apollo spacecraft shields were designed to protect astronauts from high friction temperatures (well over 2,000 degrees Fahrenheit) when re-entering the Earth's atmosphere. It was necessary to test and verify the heat shield materials on Earth before space flight. After exhaustive research and testing, NASA decided to use plasma heating as a heat source. This technique involves passing a strong electric current through a rarefied gas to create a plasma (ionized gas) that produces an intensely hot flame. Although NASA did not invent the concept, its work expanded the market for commercial plasma heating systems. One company, Plasma Technology Corporation (PTC), was founded by a member of the team that developed the Re-entry Heating Simulator at Ames Research Center (ARC). Dr. Camacho, President of PTC, believes the technology has significant environmental applications. These include toxic waste disposal, hydrocarbon, decomposition, medical waste disposal, asbestos waste destruction, and chemical and radioactive waste disposal.

Source record↗
NASA NTRS2006

Patches for Repairing Ceramics and Ceramic-Matrix Composites

Patches consisting mostly of ceramic fabrics impregnated with partially cured polymers and ceramic particles are being developed as means of repairing ceramics and ceramic-matrix composites (CMCs) that must withstand temperatures above the melting points of refractory metal alloys. These patches were conceived for use by space-suited, space-walking astronauts in repairing damaged space-shuttle leading edges: as such, these patches could be applied in the field, in relatively simple procedures, and with minimal requirements for specialized tools. These design characteristics also make the patches useful for repairing ceramics and CMCs in terrestrial settings. In a typical patch as supplied to an astronaut or repair technician, the polymer would be in a tacky condition, denoted as an A stage, produced by partial polymerization of a monomeric liquid. The patch would be pressed against the ceramic or CMC object to be repaired, relying on the tackiness for temporary adhesion. The patch would then be bonded to the workpiece and cured by using a portable device to heat the polymer to a curing temperature above ambient temperature but well below the maximum operating temperature to which the workpiece is expected to be exposed. The patch would subsequently become pyrolized to a ceramic/glass condition upon initial exposure to the high operating temperature. In the original space-shuttle application, this exposure would be Earth-atmosphere-reentry heating to about 3,000 F (about 1,600 C). Patch formulations for space-shuttle applications include SiC and ZrO2 fabrics, a commercial SiC-based pre-ceramic polymer, and suitable proportions of both SiC and ZrO2 particles having sizes of the order of 1 m. These formulations have been tailored for the space-shuttle leading-edge material, atmospheric composition, and reentry temperature profile so as to enable repairs to survive re-entry heating with expected margin. Other formulations could be tailored for specific terrestrial applications.

Hogenson, Peter A.↗
NASA NTRS2005

Technical Consultation of the Hubble Space Telescope (HST) Nickel Hydrogen (NiH2) Battery Charge Capacity Prediction

The purpose of the GSFC position paper is to identify critical HST milestone dates for continued science studies followed by the attachment of a re-entry module or a robotic servicing mission. The paper examines the viability of the HST with respect to the NiH2 continued battery charge capacity. In the course of the assessment, it was recognized that the HST battery thermal control system has an average heat dissipation limitation of 30 W per bay per orbit cycle. This thermal constraint will continue to govern options for battery capacity maintenance. In addition, the HST usage represents the longest exposure ofNiH2 batteries to Low Earth Orbit (LEO) at the current level of Depth of Discharge (DOD). Finally, the current battery life is at the limit predicted by the manufacturer, Eaglepicher. Therefore, given these factors, the potential exists that the HST battery capacities could radically degrade at any point. Given this caveat on any life extrapolations, the conservative model proposed in the GSFC position paper was viewed by the NESC as having several technical assumptions such as limited utilization of flight battery capacity data, the susceptibility of the proposed prediction method to large variations when supplemented with additional information, and the failure to qualitatively or quantitatively assess life prediction sensitivities. The NESC conducted an independent evaluation of the supporting information and assumptions to generate the predictions for battery capacity loss and practicality of on-orbit battery conditioning.

Gentz, Steven J.↗
NASA NTRS1955

A comparative analysis of the performance of long-range hypervelocity vehicles

Long-range hypervelocity vehicles are studied in terms of their motion in powered flight, and their motion and aerodynamic heating in unpowered flight. Powered flight is analyzed for an idealized propulsion system which rather closely approaches present-day rocket motors. Unpowered flight is characterized by a return to earth along a ballistic, skip, or glide trajectory. Only those trajectories are treated which yield the maximum range for a given velocity at the end of powered flight. Aerodynamic heating is treated in a manner similar to that employed previously by the senior authors is studying ballistic missiles (NACA RM A53D28), with the exception that radiant as well as convective heat transfer is considered in connection with glide and skip vehicles. The ballistic vehicle is found to be the least efficient of the several types studied in the sense that it generally requires the highest velocity at the end of powered flight in order to attain a given range. This disadvantage may be offset, however, by reducing convective heat transfer to the re-entry body through the artifice of increasing pressure drag in relation to friction drag - that is, by using a blunt body. Thus the kinetic energy required by the vehicle at the end of powered flight may be reduced by minimizing the mass of coolant material involved.

Eggers, Alfred J , Jr↗
NASA NTRS2001

Re-Entry Survivability Risk Assessment of the Extreme Ultraviolet Explorer (EUVE)

A reentry analysis of the Extreme Ultraviolet Explorer (EUVE) spacecraft was performed using the Object Reentry Survival Analysis Tool (ORSAT) - Version 5.0. The analysis was done in response to a request by NASA Headquarters and Goddard Space Flight Center (GSFC) after a preliminary assessment using the NASA Johnson Space Center Debris Assessment Software (DAS) - Version 1.0 had shown that the EUVE reentry may produce a debris area greater than the limit set within the NASA Safety Standard 1740.14 guidelines. DAS predicted that an uncontrolled reentry of the EUVE spacecraft would result in a total casualty area of 12.41 sq m, which exceeds the 8 sq m limit set in the NASA standards and implies a potential human casualty risk of approximately 1 in 5300. The ORSAT model enabled a higher fidelity thermal analysis of the EUVE spacecraft, utilizing sophisticated material and thermal properties such as emissivity, heat of oxidation, thermal conductivity, and material thickness inputs, which provided a foundation for a more in depth analysis of the reentering objects. Due to the conservative nature of the DAS study, it was reasonable to run ORSAT for only the ten objects shown to survive in the original DAS analysis. The result of the ORSAT study was a reduced casualty area of only 5.95 sq m, well within NASA safety limits. With the risk to human life now acceptably low, NASA can avoid having to take mitigation measures and allow EUVE to reenter the Earth's atmosphere uncontrolled.

O'Hara, Robin E.↗
NASA NTRS1994

Ways to Increase Launch Velocities of 2-Stage Gas Guns

The amount of space debris is rapidly increasing and the debris is distributed over a wide variety of orbits. Satellites, manned space vehicles and space stations will have to pay increasing attention to the dangers of impacts with space debris. Various armoring techniques (i.e., double or triple layer armor) will have to tested extensively to determine the most effective armor per unit weight. Intersecting near-earth orbits can lead to impact velocities up to 15 km/sec. Conventional two-stage light gas guns can launch intact, controlled-shape projectiles with a density of 1.2 gm/cc and length- to-diameter ratios of 0.5-1.0 at velocities up to 8-9 km/sec. Higher velocities (10-11 km/sec) can be obtained' for very light projectiles. The higher launch velocities tend to be very severe on the high pressure coupling and barrel of the gun and lead to short component lifetimes. Clearly, the ability to raise the launch velocity of a gun (for reasonably massive projectile shapes) from 8-9 km/sec to 11-13 km/sec (or higher), without reduction of component lifetimes, would have significant benefits. This would allow much better simulation of the higher velocity debris impacts as well as better simulation of high speed re-entry into planetary atmospheres. Several techniques for increasing the launcher muzzle velocity above 8-9 km/sec have been studied using CFD simulations and appear to offer the potential for significant gains. The first technique is to use multiple compressions, instead of a single compression, in the pump tube of the light gas gun. In a sense, this is a kind of pre-heating of the gas in the pump tube; other types of pre-heating have yielded disappointing results in the past. The dynamics of the multiple compression pump tube is very different, however, from the earlier techniques, where the pump tube was typically heated ohmically before the gun 2 cycle was started. In this paper, we present CFD calculations that show that significant increases in muzzle velocity can be obtained with multiple compressions in the pump tube. With a conventional two-stage gun, an important limitation to obtaining higher velocities is friction and heat transfer to the barrel, which typically has a length- to- diameter ratio of 200-400. These viscous losses greatly reduce the effectiveness of the regions of the barrel far removed from the second stage breech. We have studied computationally the effect of adding an additional breech (or breeches) along the barrel to reduce these viscous losses. Velocity increases from 6.5 to 7.2 km/sec have been obtained using the main breech and one additional breech. In these results, both breeches were operated with hydrogen, heated electrothermally. We have also studied a gun geometry where the main breech is operated in the conventional manner, using piston compression. The additional breech is operated either with electrothermal heating or heating by using a high explosive charge in a novel geometry. The latter option provides very effective compression, heating and acceleration of the hydrogen working gas and is fully reusable. Calculations are presented which show that very substantial increases in muzzle velocity can be obtained this way, without overstressing the projectile or the 'gun. The third technique studied is to add a section of ram accelerator tube after the barrel to further accelerate the projectile. The ram accelerator used here is not the conventional premixed gas ram accelerator, but a new technique using high explosive as the energy source and pure hydrogen as the working gas in a geometry which can be made fully reusable. Preliminary results with this new rain accelerator geometry were presented and showed that stable ram accelerator drive can be established. Herein, detailed calculations axe presented which show that substantial velocity increases can be obtained using this ram accelerator technique in tandem with a conventional light gas gun.

Bogdanoff, David W.↗
NASA NTRS1958

Three-Dimensional Orbits of Earth Satellites, Including Effects of Earth Oblateness and Atmospheric Rotation

The principal purpose of the present paper is to present sets of equations which may be used for calculating complete trajectories of earth satellites from outer space to the ground under the influence of air drag and gravity, including oblateness effects, and to apply these to several examples of entry trajectories starting from a circular orbit. Equations of motion, based on an "instantaneous ellipse" technique, with polar angle as independent variable, were found suitable for automatic computation of orbits in which the trajectory consists of a number of revolutions. This method is suitable as long as the trajectory does not become nearly vertical. In the terminal phase of the trajectories, which are nearly vertical, equations of motion in spherical polar coordinates with time as the independent variable were found to be more suitable. In the first illustrative example the effects of the oblateness component of the earth's gravitational field and of atmospheric rotation were studied for equatorial orbits. The satellites were launched into circular orbits at a height of 120 miles, an altitude sufficiently high that a number of revolutions could be studied. The importance of the oblateness component of the earth's gravitational field is shown by the fact that a satellite launched at circular orbital speed, neglecting oblateness, has a perigee some 67,000 feet lower when oblateness forces are included in the equations of motion than when they are not included. Also, the loss in altitude per revolution is double that of a satellite following an orbit not subject to oblateness. The effect of atmospheric rotation on the loss of altitude per revolution was small. As might be surmised, the regression of the line of nodes as predicted by celestial mechanics is unchanged when drag is included. It is clear that the inclination of the orbital plane to the equator will be relatively unaffected by drag for no atmospheric rotation since the drag lies in the orbital plane in this case. With the inclusion of atmospheric rotation it was found that the inclination of the plane changed about one-millionth of a radian per revolution. Thus the prediction of the position of the orbital plane of an earth satellite is not complicated by the introduction of drag. The line of apsides, which without drag but with oblateness moves slowly in space, tends to move with the satellite when drag is included in the calculations. As a results, the usual linearized solutions based on oblateness alone must be basically altered when drag is included to take into account the rapid movement of the line of apsides. In the second illustrative example the final revolution was calculated to impact for a number of trajectories in an orbital plane inclined at 650 to the equator. Of particular interest is the large effect the oblateness gravitational field and atmospheric rotation can have on the impact point. For a value of CDA/m of unity, and for an initial downward angle at 80 miles altitude of 0.01 radian, such as might be utilized for manned re-entry, oblateness had an influence of about 300 miles in the impact point, and atmospheric rotation had about a 150-mile influence.

Nielsen, Jack N.↗
NASA NTRS2002

An Automated Method to Compute Orbital Re-Entry Trajectories with Heating Constraints

Determining how to properly manipulate the controls of a re-entering re-usable launch vehicle (RLV) so that it is able to safely return to Earth and land involves the solution of a two-point boundary value problem (TPBVP). This problem, which can be quite difficult, is traditionally solved on the ground prior to flight. If necessary, a nearly unlimited amount of time is available to find the "best" solution using a variety of trajectory design and optimization tools. The role of entry guidance during flight is to follow the pre-determined reference solution while correcting for any errors encountered along the way. This guidance method is both highly reliable and very efficient in terms of onboard computer resources. There is a growing interest in a style of entry guidance that places the responsibility of solving the TPBVP in the actual entry guidance flight software. Here there is very limited computer time. The powerful, but finicky, mathematical tools used by trajectory designers on the ground cannot in general be made to do the job. Nonconvergence or slow convergence can result in disaster. The challenges of designing such an algorithm are numerous and difficult. Yet the payoff (in the form of decreased operational costs and increased safety) can be substantial. This paper presents an algorithm that incorporates features of both types of guidance strategies. It takes an initial RLV orbital re-entry state and finds a trajectory that will safely transport the vehicle to a Terminal Area Energy Management (TAEM) region. During actual flight, the computed trajectory is used as the reference to be flown by a more traditional guidance method.

Zimmerman, Curtis↗
NASA NTRS2007

Ceramic Paste for Patching High-Temperature Insulation

A ceramic paste that can be applied relatively easily, either by itself or in combination with one or more layer(s) of high-temperature ceramic fabrics, such as silicon carbide or zirconia, has been invented as a means of patching cracks or holes in the reinforced carbon-carbon forward surfaces of a space shuttle in orbit before returning to Earth. The paste or the paste/fabric combination could also be used to repair rocket-motor combustion chambers, and could be used on Earth to patch similar high-temperature structures. The specified chemical composition of the paste admits of a number of variations, and the exact proportions of its constituents are proprietary. In general, the paste consists of (1) silicon carbide, possibly with addition of (2) hafnium carbide, zirconium carbide, zirconium boride, silicon tetraboride, silicon hexaboride, or other metal carbides or oxides blended with (3) a silazane-based polymer. Because the paste is viscous and sticky at normal terrestrial and outer-space ambient temperatures, high-temperature ceramic fabrics such as silicon carbide or zirconia fabric impregnated with the paste (or the paste alone) sticks to the damaged surface to which it is applied. Once the patch has been applied, it is smoothed to minimize edge steps as required [forward-facing edge steps must be < or equal to 0.030 in. (< or equal to 0.76 mm) in the original intended space-shuttle application]. The patch is then heated to a curing temperature thereby converting it from a flexible material to a hard, tough material. The curing temperature is 375 to 450 F (approx.190 to 230 C). In torch tests and arc-jet tests, the cured paste was found to be capable of withstanding a temperature of 3,500 F (approx. 1,900 C) for 15 minutes. As such, the material appears to satisfy the requirement, in the original space-shuttle application, to withstand re-entry temperatures of approx.3,000 F (approx. 1,600 C).

Adam, Steven J.↗
NASA NTRS1995

A Simple Space Station Rescue Vehicle

Early in the development of the Space Station it was determined that there is a need to have a vehicle which could be used in the event that the Space Station crew need to quickly depart and return to Earth when the Space Shuttle is not available. Unplanned return missions might occur because of a medical emergency, a major Space Station failure, or if there is a long-term interruption in the delivery of logistics to the Station. The rescue vehicle ms envisioned as a simple capsule-type spacecraft which would be maintained in a dormant state at the Station for several years and be quickly activated by the crew when needed. During the assembly phase for the International Space Station, unplanned return missions will be performed by the Russian Soyuz vehicle, which can return up to three people. When the Station assembly is complete there will be a need for rescue capability for up to six people. This need might be met by an additional Soyuz vehicle or by a new vehicle which might come from a variety of sources. This paper describes one candidate concept for a Space Station rescue vehicle. The proposed rescue vehicle design has the blunt-cone shape of the Apollo command module but with a larger diameter. The rescue vehicle would be delivered to the Station in the payload bay of the Space Shuttle. The spacecraft design can accommodate six to eight people for a one-day return mission. All of the systems for the mission including deorbit propulsion are contained within the conical spacecraft and so there is no separate service module. The use of the proven Apollo re-entry shape would greatly reduce the time and cost for development and testing. Other aspects of the design are also intended to minimize development cost and simplify operations. This paper will summarize the evolution of rescue vehicle concepts, the functional requirements for a rescue vehicle, and describe the proposed design.

Petro, Andrew↗
NASA NTRS1994

Nonequilibrium radiation and chemistry models for aerocapture vehicle flowfields

The primary accomplishments of the project were as follows: (1) From an overall standpoint, the primary accomplishment of this research was the development of a complete gasdynamic-radiatively coupled nonequilibrium viscous shock layer solution method for axisymmetric blunt bodies. This method can be used for rapid engineering modeling of nonequilibrium re-entry flowfields over a wide range of conditions. (2) Another significant accomplishment was the development of an air radiation model that included local thermodynamic nonequilibrium (LTNE) phenomena. (3) As part of this research, three electron-electronic energy models were developed. The first was a quasi-equilibrium electron (QEE) model which determined an effective free electron temperature and assumed that the electronic states were in equilibrium with the free electrons. The second was a quasi-equilibrium electron-electronic (QEEE) model which computed an effective electron-electronic temperature. The third model was a full electron-electronic (FEE) differential equation model which included convective, collisional, viscous, conductive, vibrational coupling, and chemical effects on electron-electronic energy. (4) Since vibration-dissociation coupling phenomena as well as vibrational thermal nonequilibrium phenomena are important in the nonequilibrium zone behind a shock front, a vibrational energy and vibration-dissociation coupling model was developed and included in the flowfield model. This model was a modified coupled vibrational dissociation vibrational (MCVDV) model and also included electron-vibrational coupling. (5) Another accomplishment of the project was the usage of the developed models to investigate radiative heating. (6) A multi-component diffusion model which properly models the multi-component nature of diffusion in complex gas mixtures such as air, was developed and incorporated into the blunt body model. (7) A model was developed to predict the magnitude and characteristics of the shock wave precursor ahead of vehicles entering the Earth's atmosphere. (8) Since considerable data exists for radiating nonequilibrium flow behind normal shock waves, a normal shock wave version of the blunt body code was developed. (9) By comparing predictions from the models and codes with available normal shock data and the flight data of Fire II, it is believed that the developed flowfield and nonequilibrium radiation models have been essentially validated for engineering applications.

Carlson, Leland A.↗
NASA NTRS1999

Exploring Mars: the Ares Payload Service (APS)

In last year's Mars Society convention we introduced the results of five years of studies of space launch capability for the second millennium. We concluded that Single Stage to Orbit (SSTO) vehicles such as the Delta Clipper X33, and X34 cannot make it to orbit from the Earth's surface. Whether taking off vertically or horizontally or landing vertically or horizontally, the rocket equations, the performance of available fuels, and the realities of the weight and strength of materials leave no margin for payload. The promised savings from SSTO systems are illusory. However, a configuration that is able to deliver useful payload to orbit is the Single step to Orbit, SsTO, a rocket plane that is released fully fueled, from 35,000 to 40,000 feet altitude. Three approaches have been proposed. The Hot'l and Molnya Corporation designs carry the fueled rocket plane to altitude on the back of a carrier aircraft. In this design the carrier aircraft is Russia's Antonov 225 the world's largest cargo plane. The rocket plane is a modified version of the Buran, Russia's own space shuttle. Another configuration is Kelly Aviation's concept in which the fully fueled rocket plane is towed to altitude by the cargo plane and then released. A third approach is based on the early "X" planes, which were dropped from the belly of the carrier plane. While the rocket equations indicate that these three concepts can deliver useful payloads, the Stanford review found significant advantages to the approach of Pioneer Rocket, in which the rocket plane flies up to the carrier plane with conventional jet engines, docks, and then loads on the oxidizer for the flight to orbit. This architecture has more reasonable abort modes in case of system failure in either aircraft and can deliver a larger final payload to orbit for a given sized carrier. The Stanford recommendation is that the carrier aircraft be the Antonov 225. A design based on this was presented in a report last year. Refinements to the design notably an improved re-entry cooling system and fueling stability analysis were done this year. More technical detail and a proposed international consortium to develop the SSTO is presented in another session of this year's Mars convention. We believe that there will be no human exploration of Mars based on the Shuttle or Expendable launch vehicles, and no resources available except for a cooperative international program. However, just as the world is learning to cooperate in peacekeeping, we hold out the hope that similar cooperation will develop for Mars exploration. With that in mind, this year we asked the question- "How will the human mission get to Mars if it has to use the SsTO for transportation?"

Bowen, Justin↗
NASA NTRS1993

The role of visual context in manual target localization

During space flight and immediately after return to the 1-g environment of earth, astronauts experience perceptual and sensory-motor disturbances. These changes result from adaptation of the astronaut to the microgravity environment of space. During space flight, sensory information from the eyes, limbs, and vestibular organs is reinterpreted by the central nervous system in order to produce appropriate body movements in the microgravity. This adaptation takes several days to develop. Upon return to earth, the changes in the sensory-motor system are no longer appropriate to a 1-g environment. Over several days, the astronaut must re-adapt to the terrestrial environment. Alterations in sensory-motor function may affect eye-head-hand coordination and, thus, the crewmember's ability to manually locate objects in extrapersonal space. Previous reports have demonstrated that crewmembers have difficulty in estimating joint and limb position and in pointing to memorized target positions on orbit and immediately postflight. The ability to point at or reach toward an object or perform other manual tasks is essential for safe Shuttle operation and may be compromised particularly during re-entry and landing sequences and during possible emergency egress from the Shuttle. An understanding of eye-head-hand coordination and the changes produced during space flight is necessary to develop effective countermeasures. This summer's project formed part of the study of the sensory cues use in the manual localization of objects.

Barry, Susan R.↗
NASA NTRS2005

Diode Laser Sensors for Arc-Jet Characterization

The development and application of tunable diode laser (TDL) absorption sensors to monitor the health and operating conditions in the large-scale 60 MW arc-heated- plasma wind-tunnel at NASA Ames Research Center is reported. The interactive heating facility (THF) produces re-entry flow conditions by expanding the gas heated in a constricted plasma arc-heater to flow at high velocity over a model located in a test cabin. This facility provides the conditions needed to test thermal protective systems for spacecraft re-entering the earth s atmosphere. TDL sensors are developed to monitor gas flows in both the high-temperature constricted flow and the supersonic expansion flow into test cabin. These sensors utilize wavelength-tuned diode lasers to measure absorption transitions of atomic oxygen near 777.2 nm, atomic nitrogen near 856.8 nm, and atomic copper near 793.3 nm. The oxygen and nitrogen sensors measure the population density in exited electronic states of these atoms. The measurements combined with the assumption of local thermal and chemical equilibrium yield gas temperature (typically near 7,000K). The nitrogen and oxygen population temperatures are redundant, and their close agreement provides an important test of the local thermal equilibrium assumption. These temperature sensors provide time-resolved monitors of the operating conditions of the arc-heater and can be used to verify and control the test conditions. An additional TDL sensor was developed to monitor the copper concentration in the arc-heater flow yielding values as high as 13 ppm. Measurements of copper in the flow can identify flow conditions with unacceptably rapid electrode erosion, and hence this sensor provides valuable information needed to schedule maintenance to avoid costly arc-heater failure. TDL sensors were also developed for measurements in the test cabin, where absorption measurements of the populations of argon and molecular nitrogen in excited metastable electronic states established that the number density of these excited species is much lower than estimated using frozen-chemistry approximations. This key finding suggests that in the post-expansion region there is not a significant energy sequestration in electronically excited species. Finally, TDL measurements of atomic potassium seeded into the test cabin flow were used to directly measure the static temperature of the test gas. The results of this study illustrate the high potential of time-resolved TDL measurements for routine and economical sensing of arc-heater health (gas temperature and electrode erosion) as well as the time-resolved test-cabin-flow conditions in front of the model.

Hanson, Ronald K.↗
NASA NTRS2006

Entry, Descent, and Landing Operations Analysis for the Stardust Re-Entry Capsule

On the morning of January 15, 2006, the Stardust capsule successfully landed at the Utah Test and Training range in northwest Utah returning cometary samples from the comet Wild-2. An overview of the entry, descent, and landing (EDL) trajectory analysis that was performed for targeting during the Stardust Mission Navigation Operations Phase upon final approach to Earth is described. In addition, how the predicted landing location and the resulting overall 99 percentile landing footprint ellipse obtained from a Monte Carlo analysis changed over the final days and hours prior to entry is also presented. The navigation and EDL operations effort accurately delivered the entry capsule to the desired landing site. The final landing location was 8.1 km from the target, which was well within the allowable landing area.

Desao, Prasun N.↗