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Subcentimeter orbital debris is currently undetectable using ground-based radar and optical methods. However, pits in space shuttle windows produced by paint chips demonstrate that small debris can...

The vertical landing and recovery of rockets are now considered an important solution in the field of reusable launch vehicles. This paper focuses on guided atmospheric entry to reach an expected l...

Designed for the launch and subsequent on-orbit maintenance by astronauts and the Space Shuttle, the Hubble Space Telescope presented a unique and challenging set of flight operations issues during six missions between 1990 and 2009. Starting with the launch in 1990 and for the following three servicing missions, the solar arrays represented an unanticipated and significant operational challenge for the crews and the flight operations teams. Modified first-generation solar arrays were installed on Servicing Mission 1 in 1993 to correct for a jitter introduced into the telescope’s control system by the design of the original solar arrays. However, although the jitter was improved, the modified solar arrays introduced additional issues for the second servicing mission, Space Transportation System 82, due to the existence of a significant static twist. New flight operations techniques had to be developed to monitor the condition and protect the integrity of the solar arrays. It was not known until after Hubb...

The Hubble Space Telescope is unquestionably one of the most important scientific instruments ever developed. Designed for launch and subsequent on-orbit maintenance by astronauts and the space shuttle, the Hubble Space Telescope presented a unique and challenging set of flight operations issues. The complexity of the operations required a sophisticated level of mission planning and execution, particularly in robotics and extravehicular activity. Detailed planning by the flight operations team and the crew before the launch of the Hubble Space Telescope onboard the space shuttle Discovery was critical to the successful achievement of the objectives. The on-orbit execution of the mission plan, the anomalies encountered, and the resulting lessons learned from the deployment mission were important not only for preparations for the later servicing missions but also for a variety of missions, including those dedicated to the International Space Station assembly.

The solid rocket booster main flame deflector at NASA Kennedy Space Center Launch Complex 39A was instrumented to measure heat rates, pressures, and temperatures on the final three space shuttle launches. Because the solid rocket booster plume is hot and erosive, a robust tungsten piston calorimeter was developed to compliment measurements made by off-the-shelf sensors. Witness materials were installed, and their melting and erosion response to the Mach 2, 4000°F, 4 s duration plume was observed. The data show that the specification used for the design of the main flame deflector thermal protection system overpredicts heat rates by a factor of three and underpredicts pressures by a factor of two. The discovery of short-duration heating spikes that occur when aluminum oxide slag solidifies on the main flame deflector explains this heat rate overprediction. This study allows improvement of solid rocket motor launch site and test stand computational fluid dynamics models and the concomitant slag deposition h...

A = axial location of the solid rocket booster (SRB) reference point in Space Shuttle main engine (SSME) plume coordinate system, ft ax, ay, az = acceleration components along body axes (excluding components due to gravity), ft=s CA = isolated (freestream) axial force coefficient Cm = isolated (freestream) pitching moment coefficient CN = isolated (freestream) normal force coefficient Cn = isolated (freestream) yawing moment coefficient CY = isolated (freestream) side force coefficient Cl = isolated (freestream) rolling moment coefficient eA, eB = unit vectors in body A and body B, respectively Fp = plume impingement force, lb Fx, Fy, Fz = aerodynamic forces in axial, lateral, and normal directions, lb F CON A , F CON B = joint constraint force vector for body A and body B

A structural-health-monitoring system can contribute to the risk management of a structure operating under hazardous conditions.An example is thewing leading-edge impact-detection system thatmonitors the debris hazards to the Space Shuttle Orbiter’s reinforced carbon–carbon panels. Since return to flight after the Columbia accident, the system was developed and subsequently deployed on board the orbiter to detect ascent and on-orbit debris impacts, so as to support the assessment ofwing leading edge structural integrity before orbiter reentry.As structural healthmonitoring is inherently an inverse problem, the analyses involved, including those performed for this system, tend to be associated with significant uncertainty. The use of probabilistic approaches to handle the uncertainty has resulted in the successful implementation of many development and application milestones.

Recent, current, and planned NASA missions that employ blunt-body entry vehicles pose aerothermodyamic problems that challenge state-of-the-art experimental and computational methods. The issues of boundary-layer transition and turbulent heating on the heat shield have become important in the designs of both the Mars Science Laboratory and Crew Exploration Vehicle. While considerable experience in these general areas exists, that experience is mainly derived from simple geometries; e.g., sharp-cones and flat-plates, or from lifting bodies such as the Space Shuttle Orbiter. For blunt-body vehicles, application of existing data, correlations, and comparisons is questionable because an all, or mostly, subsonic flowfield is produced behind the bow shock, as compared with the supersonic (or even hypersonic) flow of other configurations. Because of the need for design and validation data for projects such asMars Science Laboratory andCrewExplorationVehicle, many new experimental studies have been conducted in the last decade to obtain detailed boundary-layer transition and turbulent heating data on this class of vehicle. In this paper, details of several of the test programs are reviewed. The laminar and turbulent data from these various test are shown to correlate in terms of edge-based Stanton andReynolds number functions. Correlations are developed from the data for transition onset and turbulent heating augmentation as functions of momentum thickness Reynolds number. These correlations can be employed as engineering-level design and analysis tools.

A piloted simulation that studied the handling qualities for a precision lunar landing task from final approach to touchdown is presented. A vehicle model based on NASA's Altair Lunar Lander was used to explore the design space around the nominal vehicle configuration to determine which combination of factors provides satisfactory pilot-vehicle performance and workload; details of the control and propulsion systems not available for that vehicle were derived from Apollo Lunar Module data. The experiment was conducted on a large motion base simulator. Eight Space Shuttle and Apollo pilot astronauts and three NASA test pilots served as evaluation pilots, providing Cooper-Harper ratings, Task Load Index ratings and qualitative comments. Each pilot flew seven combinations of control response types and three sets of displays, including two varieties of guidance and a nonguided approach. The response types included Rate Command with Attitude Hold, which was used in the original Apollo Moon landings, a Velocity Increment Command response type designed for up-and-away flight, three response types designed specifically for the vertical descent portion of the trajectory, and combinations of these. It was found that Velocity Increment Command significantly improved handling qualities when compared with the baseline Apollo design, receiving predominantly Level 1 ratings. This response type could be flown with or without explicit guidance cues, something that was very difficult with the baseline design, and resulted in approximately equivalent touchdown accuracies and propellant burn as the baseline response type. The response types designed to be used exclusively in the vertical descent portion of the trajectory did not improve handling qualities.

This paper addresses different methods for determining the amount of extra propellant (flight performance reserve or FPR) that is necessary to reach orbit with a high probability of success. One approach involves assuming that the various influential parameters are independent and that the result behaves as a Gaussian. Alternatively, probabilistic models may be used to determine the vehicle and environmental models that will be available (estimated) for a launch day go/no go decision. High-fidelity closed-loop Monte Carlo simulation determines the amount of propellant used with each random combination of parameters that are still unknown at the time of launch. Using the results of the Monte Carlo simulation, several methods were used to calculate the FPR. The final chosen solution involves determining distributions for the pertinent outputs and running a separate Monte Carlo simulation to obtain a best estimate of the required FPR. This result differs from the result obtained using the other methods sufficiently that the higher fidelity is warranted.

Rigid polyurethane foams and rigid polyisocyanurate foams (spray-on foam insulation), like those flown on Shuttle, Delta IV, and will be flown on Ares-I and Ares-V, can gain an extraordinary amount of water when under cryogenic conditions for several hours. These foams, when exposed for eight hours to launch pad environments on one side and cryogenic temperature on the other, increase their weight from 35 to 80 percent depending on the duration of weathering or aging. This effect translates into several thousand pounds of additional weight for space vehicles at lift-off. A new cryogenic moisture uptake apparatus was designed to determine the amount of water/ice taken into the specimen under actual-use propellant loading conditions. This experimental study included the measurement of the amount of moisture uptake within different foam materials. Results of testing using both aged specimens and weathered specimens are presented. To better understand cryogenic foam insulation performance, cryogenic moisture testing is shown to be essential. The implications for future launch vehicle thermal protection system design and flight performance are discussed.

The Rigidizable Inflatable Get-Away-Special EXperiment (RIGEX) was run successfully on board STS-123 (Endeavor) in March 2008. RIGEX was built by graduate students at the Air Force Institute of Technology (AFIT) and returned there following the shuttle flight for post-flight analysis. The experiment’s objectives were to demonstrate in space the stowage, deployment, and rigidization techniques of carbon fiber composite inflatable rigidizable cylindrical booms. RIGEX was a Canister For All Payloads (CAPE) Space Shuttle cargo bay experiment designed to heat and inflate three 50.8 cm (20 in) long carbon fiber composite booms in a microgravity vacuum environment and measure both the structural characteristics and the deployment accuracy. Pressure, temperature, modal response, and position data were collected successfully on-orbit and are compared here to ground test data. This research is intended to help demonstrate the feasibility of using lightweight and low stowage volume (high packaging ratio) inflatable/rigidizable space structures for space mission applications.

T HE shuttle ionospheric modification with pulsed localized exhaust (SIMPLEX) experiments conducted by the Naval Research Laboratory (NRL) were designed to 1) assess the effects of rocket exhaust interactions in the ionosphere and 2) mimic large ionospheric disturbances that occur naturally [1–5]. These experiments use space shuttle orbital maneuver subsystem (OMS) engines to inject exhaust over ground radar sites. The shuttle exhaust provides a high-speed neutral gas that streams through the ambient plasma of the ionosphere. The neutral exhaust molecules exchange chargewith ambient O ions in the ionosphere. This interaction between the plasma and neutrals results in the formation of ion-ring and beam velocity distributions of plasma particles in the ionosphere. During the SIMPLEX experiments, these distributions are studied with ground radars using incoherent scatter of radio waves from electrons in the ionosphere. To date, six SIMPLEX burns [1–5] have been conducted over incoherent scatter radar sites at various locations, as shown in Table 1. In these experiments, the relative velocity between the ambient ions and injected neutrals is much faster than that which occurs naturally in any region of the ionosphere. Auroral electric fields can accelerate ions to velocities between 0.5 and 2:5 m=s [6]. The neutral injections from the space shuttle in orbit yield exhaust velocities of between 4.7 and 10:7 km=s relative to the background neutrals and ions. During the shuttle burn experiments, the highspeed neutrals chemically react with the stationary plasma. During auroral plasma convection, however, the high-speed ambient O ions and electrons in the ionosphere (accelerated to a high velocity by auroral electric fields) chemically react with stationary neutral molecules or atoms [5,6]. Consequently, the primary difference between high-speed plasma convection and space shuttle OMS burns is the reference frame for the relative ion and neutral motion [5]; both phenomena result in the formation of ring-beam ion velocity distributions. The SIMPLEX experiments were designed to reproduce the naturally occurring ion-ring distributions, which can create ionospheric instabilities leading to regions of plasma turbulence. In addition, the experiments also provide a unique way to examine gas-phase physical and chemical phenomena in the hypersonic and hyperthermal energy regime, which is relatively unstudied and difficult to reproduce in a laboratory [7]. The first step to modeling the ionospheric interactions of space shuttle exhaust is to describe the neutral flow from the exit plane of the shuttle OMSengines into the expanse of the upper atmosphere. In this paper, we use the direct simulation Monte–Carlo (DSMC) method [8] to simulate the shuttle burn and to study the interaction between the shuttle exhaust and the neutral species of the background atmosphere. In the following sections, we present simulation results showing the time evolution of the shuttle plume and background, and we discuss the effect of the altitude-dependence of the background atmosphere.

The Air Force Institute of Technology (AFIT) developed the Rigidizable Inflatable GetAway-Special Experiment (RIGEX) to demonstrate the feasibility of using rigidizable, inflatable, technology for space structures. RIGEX autonomously deployed and captured the structural characteristics of three rigidizable, inflatable tubes while in the space environment. The three identical tubes were manufactured by L’Garde Inc. and were made from carbon fibers in a thermoplastic resin. RIGEX was designed to fly in the DoD’s Space Test Program’s Canister for All Payload Ejections (CAPE) container within the Space Shuttle Orbiter's payload bay. RIGEX completed its mission on March 26, 2008 after being in orbit for 16 days. This paper summarizes the science and motivation behind RIGEX's inflatable rigidizable structure research and provides an overview of the experiment design. It also details the experiment build and flight qualification testing of the RIGEX/CAPE payload, including modifications and design changes made to the experiment during the build and test phases. Results from flight qualification testing are presented which were used to validate RIGEX’s structural integrity, functionality, and space compatibility. These test results were compiled into the payload’s Acceptance Data Package which was provided to the National Air and Space Association (NASA). RIGEX was integrated and flown under the direction of DoD’s Space Test Program.

This paper explores mission scenarios using existing evolved expendable launch vehicles for delivering, on an interim basis, the crew exploration vehicle Orion to the International Space Station. The use of existing commercial launchers is proposed to narrow the International Space Station service gap from the time the space shuttle is deserviced until the new Ares I launch vehicle is crew rated and operational. Here, three launch options are evaluated:1)theAtlasVheavy-liftvehicle,2)theDeltaIVheavy-lift vehicle,and3)theDeltaIVwiththreecommoncore boosters (as a first stage, with the Orion acting as the second stage). Configurations 1 and 3 require significant impulse from the Orion’s service-module engine to achieve the final orbit. Configuration 2 launches the Orion as a passivepayload,withoutrelianceonanyimpulse fromtheservicemodule.Allthreeconfigurations reservesufficient service-module impulse for deorbit and reentry. Detailed simulation results, concepts of operation, and mission timelines are presented for each configuration. Mission feasibility is demonstrated for all three configurations. The final configuration has the advantage of eliminating failure paths and requiring human rating only for the commoncore booster on the Delta IV system. Finally, reliability- and development-cost assessments are presented and compared with the Ares I.

NASA is developing a newgeneration of spacecraft to replace the Space Shuttle and return astronauts to themoon. These spacecraft will have a manual control capability for several mission tasks, and the ease and precision with which pilots can execute these tasks will have an important effect on mission risk and training costs. A simulation evaluated the handling qualities of a generic space vehicle based on dynamics similar to one of these spacecraft, NASA’s Crew Exploration Vehicle, during the last segment of the docking task with a space station. This handling qualities evaluation looked at fourdifferent translational control systems, called response types, thatmappilot inputs to thruster firings in a way that gives predictable and useful vehicle responses. These response types were flownwith three levels of translation-into-rotation dynamic coupling arising from a longitudinal offset between the reaction control system thrusters and the vehicle’s center of mass. The results indicate that greater translation-into-rotation coupling is strongly correlated with degraded handling qualities, but that different response types do not have a major effect on pilot workload, final docking performance, or handling qualities.

An experimental wind-tunnel program is being conducted in support of a NASA wide effort to develop a Space Shuttle replacement and to support the agency’s long-term objective of returning to the moon andMars. This paper documents experimentalmeasurementsmade on several scaled ceramic heat transfermodels of the proposedProject Orion Crew Exploration Vehicle. The experimental data highlighted in this paper will be used to assess numerical tools thatwill be used to generate theflight aerothermodynamic database. Global heat transfer images and centerline heat transfer distributions were obtained over a range of freestreamReynolds numbers and angles of attackwith the phosphor thermography technique. Temperature dataweremeasured on the forebody and afterbody andwere used to compute the heating on the vehicle as well as the boundary-layer state on the forebody surface. Several model support configurations were assessed to minimize potential aftbody support interference. Although naturally fully developed turbulent levels were not obtained on the forebody, the use of boundary-layer trips generated fully developed turbulent flow. Laminar and turbulent computational results are shown to be in good agreement with the data. In addition, the ability of the global phosphor thermography method to provide quantitative heating measurements in the low-temperature environment of the capsule base regionwas assessed and the lack of significant model support hardware influence on heating was shown.

A series of 22 primary reaction-control-system engine attitude-control firings were observed from the Maui Space Surveillance Site during the space shuttle STS-115 mission. The firings occurred during a pass over Maui on 19 September 2006 during which the orbiter was in sunlight and the observatory was in darkness. The observed attitude maneuvers maintained the orbiter in an orientation in which its long axis was aligned with the line of sight from the observatory. This ensured that the thrust vectors of all the observed engine firings were perpendicular to the line of sight, providing an optimal side-on observation of the exhaust. The firings ranged between 80 and 320 ms in duration and involved 2 or 3 engines for pitch, roll and yaw adjustments. A 0.328 deg field-of-view acquisition scope of the 3.6 m telescope of the Advanced Electro-Optical System provided unfiltered imagery in the near-ultraviolet visible spectral region. The most interesting white-light features were transients, one observed at engine start up and two at shutdown. The analysis of the transient speeds reveals that the startup transient consists of either unburned propellant droplets or higher-pressure gas evaporated from droplets and that the shutdown transients are attributable to a slightly staggered release of unburned oxidizer and fuel, respectively. The first (oxidizer) shutdown transient is the brightest feature, for which an intensity evolution analysis is conducted. The analysis of the ground-based data is fully consistent with spectral features attributable to primary reaction-control-system engine transients observed in previous measurements from the space shuttle bay using an imager spectrograph.

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.

The military is pursuing a low-cost, responsive reusable launch vehicle that can rapidly place payloads in orbit in response to national defense requirements. The reusable launch vehicle is currently in the early stages of its design phase, when it is critical to assess design alternatives in terms of their operational capabilities as well as their life cycle support requirements. The U.S. Air Force Research Laboratory developed the space access vehicles mission and operations simulation (SAVMOS)model suite to assist in evaluating themission cycle of the reusable launch vehicle. Part of the functionality of SAVMOS is the maintenance, integration, and launch pad operations simulation and test (MILEPOST)model.MILEPOST allows the user to evaluate candidate designs and perform tradeoff studies for the impact of design characteristics on regeneration time and support personnel requirements. We discussMILEPOST development and present insights on process time and workforce sizing for several design alternatives.

The International Space Station will require a continuous supply of replacement parts for ongoing maintenance and repair after the planned retirement of the Space Shuttle in 2010. These parts are existing line-replaceable items collectively called Orbital Replacement Units, and include heavy and oversized items such as Control Moment Gyroscopes and stowed radiator arrays originally intended for delivery aboard the Space Shuttle. Current resupply spacecraft have limited to no capability to deliver these external logistics. In support of NASA's Exploration Systems Architecture Study, a team at Langley Research Center designed an Unpressurized Cargo Delivery Vehicle to deliver bulk cargo to the Space Station. The Unpressurized Cargo Delivery Vehicle was required to deliver at least 13,200 lbs of cargo mounted on at least 18 Flight Releasable Attachment Mechanisms. The Crew Launch Vehicle design recommended in the Exploration Systems Architecture Study would be used to launch one annual resupply flight to the International Space Station. The baseline vehicle design developed here has a cargo capacity of 16,000 lbs mounted on up to 20 Flight Releasable Attachment Mechanisms. Major vehicle components are a 5.5m-diameter cargo module containing two detachable cargo pallets with the payload, a Service Module to provide propulsion and power, and an aerodynamic nose cone. To reduce cost and risk, the Service Module is identical to the one used for the Crew Exploration Vehicle design.

Sections of the Space Shuttle External Tank Liquid Oxygen (LO2) and Liquid Hydrogen (LH2) cable trays are shielded from potentially damaging airloads with foam Protuberance Aerodynamic Load (PAL) Ramps. Flight standard design LO2 and LH2 cable tray sections were tested with and without PAL Ramp models in the United States Air Force Arnold Engineering Development Center s (AEDC) 16T transonic wind tunnel to obtain experimental data on the aeroelastic stability and response characteristics of the trays and as part of the larger effort to determine whether the PAL ramps can be safely modified or removed. Computational Fluid Dynamic simulations of the full-stack shuttle launch configuration were used to investigate the flow characeristics around and under the cable trays without the protective PAL ramps and to define maximum crossflow Mach numbers and dynamic pressures experienced during launch. These crossflow conditions were used to establish wind tunnel test conditions which also included conservative margins. For all of the conditions and configurations tested, no aeroelastic instabilities or unacceptable dynamic response levels were encountered and no visible structural damage was experienced by any of the tested cable tray sections. Based upon this aeroelastic characterization test, three potentially acceptable alternatives are available for the LO2 cable tray PAL Ramps: Mini-Ramps, Tray Fences, or No Ramps. All configurations were tested to maximum conditions, except the LH2 trays at -15 deg. crossflow angle. This exception is the only caveat preventing the proposal of acceptable alternative configurations for the LH2 trays as well. Structural assessment of all tray loads and tray response measurements from launches following the Shuttle Return To Flight with the existing PAL Ramps will determine the acceptability of these PAL Ramp alternatives.

Twomethods to infer freestream densities from in-flightmeasurements of pitot pressure and flight velocity during reentry are presented that focus on the minimization of uncertainties due to high-temperature real-gas effects and atmospheric densityfluctuation. A numerical approach leads to a curve-fit function that yields the ratio of the pitot to the dynamic pressure pt2=q1 for velocities between 0.5 and 10 km=s and altitudes up to 90 km. An analytically derived correlation is also provided. Both techniques account for equilibrium real-gas effects thus achieving very high accuracies. Independently of potential atmospheric fluctuation, remaining errors are less than 1% for almost the entire spectrum and less than 0.3% for typical lifting reentry.

Lift and drag measurements have been analyzed for subsonic flight conditions for seven blunt-based reentry-type vehicles. Five of the vehicles are lifting bodies (M2-F1, M2-F2, HL-10, X-24A, and X-24B) and two are wing-body configurations (the X-15 and the Space Shuttle Enterprise). Base pressure measurements indicate that the base drag for full-scale vehicles is approximately three times greater than predicted by Hoerner's equation for three-dimensional bodies. Base drag and forebody drag combine to provide an optimal overall minimum drag (a drag "bucket") for a given configuration. The magnitude of this optimal drag, as well as the associated forebody drag, is dependent on the ratio of base area to vehicle wetted area. Counter-intuitively, the flight-determined optimal minimum drag does not occur at the point of minimum forebody drag, but at a higher forebody drag value. It was also found that the chosen definition for reference area for lift parameters should include the projection of planform area ahead of the wing trailing edge (i.e., forebody plus wing). Results are assembled collectively to provide a greater understanding of this class of vehicles than would occur by considering them individually.

DOI: 10.2514/1.23082 This paper describes the analysis of localized catalytic heating effects to the U.S. Space Shuttle Orbiter thermal protectionsystem.Theanalysisappliestothehigh-temperaturereusablesurfaceinsulationonthelowerfuselageand wing acreage, as well as the reinforced carbon–carbon on the nose cap, chin panel, and wing leading edge. The objective was to use a modified two-layer approach to predict the catalytic heating effects on the Orbiter windward thermal protection system assuming localized highly catalytic or fully catalytic surfaces. The method incorporated the boundary layer integral matrix procedure–kinetic code with streamline inputs from viscous Navier–Stokes solutions to produce heating rates for localized fully catalytic and highly catalytic surfaces as well as for nominal partially catalytic surfaces (either reinforced carbon–carbon or reaction cured glass) with temperature-dependent recombination coefficients. The highly catalytic heating results showed very good correlation with Orbiter experiments STS-2, -3, and -5 centerline and STS-5 wing flight data. Recommended catalytic heating factors were generated for use in future shuttle missions to perform quick-time atmospheric reentry analysis of damaged or repaired thermal protection system areas. The catalytic factors are presented along streamlines and as a function of stagnation enthalpy for use with arbitrary shuttle trajectories.

Space Shuttle rendezvous missions present unique challenges that were not fully recognized when the Shuttle was designed. Rendezvous targets could be passive (i.e., no lights or transponders), and not designed to facilitate Shuttle rendezvous, proximity operations, and retrieval. Shuttle reaction control system jet plume impingement on target spacecraft presented induced dynamics, structural loading, and contamination concerns. These issues, along with limited reaction control system propellant in the Shuttle nose, drove a change from the legacy Gemini/Apollo coelliptic profile to a stable orbit profile, and the development of new proximity operations techniques. Multiple scientific and on-orbit servicing missions, and crew exchange, assembly and replenishment flights to Mir and to the International Space Station drove further profile and piloting technique changes. These changes included new proximity operations, relative navigation sensors, and new computer generated piloting cues. However, the Shuttle's baseline rendezvous navigation system has not required modification to place the Shuttle at the proximity operations initiation point for all rendezvous missions flown.

A reanalysis of discrete roughness boundary-layer transition data using a consistent computational method for comparison to other published results has been completed. The primary objective of this effort was to investigate the influence of the computational approach on the resulting transition correlation. The experimental results were previously obtained on Space Shuttle Orbiter models in the NASA Langley Research Center 20-Inch Mach 6 Air Tunnel. The phosphor thermography system was used to monitor the status of the boundary layer via global heat- transfer images of the orbiter windward surface. The existing roughness transition database included a variation in the size and location of discrete roughness trips along the centerline of 0.0075-scale models at an angle of attack of 40 deg. Various correlative approaches were attempted, with the roughness transition correlations based on edge properties providing the most reliable results. When a consistent computational method is used to compute edge conditions, transition data sets for different moderately blunt configurations at several angles of attack are shown to collapse to a well-behaved correlation. The shuttle experimental dataset presented herein, therefore, can be used to calibrate the preferred computational method of the end user for use in the future designs of the next-generation space access vehicles.

Optically sensitive surfaces on the International Space Station (ISS) can be damaged (or eroded/pitted) when impacted by high-velocity particles from unburned liquid propellant present in bipropellant thruster plumes. Surfaces with thin optical coatings, such as solar arrays and radiators, are of primary concern. Thruster plume-induced erosion/pitting of sensitive surfaces has been observed on space shuttle flight experiments. The Boeing ISS Environments Team in Houston has developed an approach to modeling thruster plume-induced erosion/pitting of ISS surface materials. The Boeing team has conducted analyses simulating bipropellant thruster particles impacting sensitive ISS surfaces for various assembly stages. Thruster firings for ISS reboost/attitude control, as well as visiting vehicle thruster firings during approach or separation to ISS docking ports, were simulated. The results of these analyses show that particle impingement angle greatly affects surface damage, with normal impacts being the most severe. Particles with highly oblique impact angles (∼75 deg off normal), however, will essentially skid off surfaces without causing any erosion/pitting. A mitigation technique has been developed to prevent plume erosion/pitting of solar array coatings. Before a thruster-firing event, solar arrays may be rotated to a preestablished position that will eliminate plume particle impact damage to the surface. The preestablished positions are defined based on the geometry of the ISS thrusters relative to the solar array panels to ensure that plume particles will impinge at highly oblique angles (greater than 75 deg off normal). Upcoming ISS milestones will introduce new sensitive surfaces and thrusters, making 2005 a critical year for establishing operational constraints to mitigate thruster plume erosion. Some of these milestones include the space shuttle return to flight, the deployment of new ISS solar arrays, and the maiden voyage of ESA's automated transfer vehicle. Operational constraints for plume erosion mitigation are being coordinated with other solar array operational constraints such as power, thermal, and plume-induced structural loads. An integrated operational solution is being implemented to support the ISS assembly flight sequence. This paper will discuss plume erosion analyses and the implementation of operational mitigation as well as ongoing testing to better characterize plume erosion effects.

Spacecraft operating in low earth orbit face a significant orbital debris impact hazard. Of particular concern, in the case of the Space Shuttle, are impacts on critical components of the thermal protection system. Recent research has formulated a new material model of reinforced carbon-carbon, for use in the analysis of hypervelocity impact effects on the Space Shuttle wing leading edge. The material model has been validated in simulations of published impact experiments and applied to model orbital debris impacts at velocities beyond the range of current experimental methods. The results suggest that momentum scaling may be used to extrapolate the available experimental data base, in order to predict the size of wing leading edge perforations at impact velocities as high as 13 km/s.

A wind-tunnel investigation of the effectiveness of an aerodynamic yaw controller mounted on the lower surface of a Space Shuttle Orbiter model body flap was conducted in the Langley 31-Inch Mach 10 tunnel. The controller, consisting of a 60-deg delta fin mounted perpendicular to the body-flap lower surface and yawed 30 deg to the freestream direction, was tested at orbiter angles of attack from 20 to 40 deg at zero sideslip for a Reynolds number based on the wing mean aerodynamic chord of 0.66 x 10 6 . The aerodynamic control characteristics are presented with an analysis of the effectiveness of the controller in making a bank maneuver for Mach 18 flight conditions. The analysis shows that the controller is as effective as the reaction control system in initiating the bank maneuver. These results warrant further studies to determine the aerodynamic heating on the control surfaces and the effects of controller hinge-line cant angle and body-flap deflection on the controller effectiveness.

There are four basic thermal materials used on the space shuttle Columbia, Reinforced-Carbon-Carbon (RCC), Low- and High Temperature Reusable Surface Insulation tiles (LRSI and HRSI, respectively), and Felt Reusable Surface Insulation (FRSI) blanket 1 . The purpose of this work was to obtain the mechanical behavior of the materials involved in the modeling of the foam impact scenario for the shuttle accident investigation. Experiments were designed to provide the modeling parameters of HRSI tiles and RCC as well as to give insight into the failure phenomena under different loading conditions. For each material the results are presented as a function of orientation and applied strain rate. The effect of aging on the RCC material properties is also discussed.

A series of three-dimensional simulations has been performed to investigate analytically the effect of insulating foam impacts on ceramic tile and reinforced carbon-carbon components of the space shuttle thermal protection system. The simulations employed a hybrid particle finite element method and a parallel code developed for use in spacecraft design applications. The conclusions suggested by the numerical study are in general consistent with experiment. The results emphasize the need for additional material testing work on the dynamic mechanical response of thermal-protection-system materials and additional impact experiments for use in validating computational models of impact effects.

The global positioning system (GPS) receiver used for navigation on the space shuttle exhibits range rate noise that appears to result from scintillation of the satellite signals by irregularities in ionospheric plasma. The noise events cluster in geographic regions previously identified as being susceptible to instability and disturbed ionospheric conditions. These mechanisms are reviewed in the context of the GPS observations. Range-rate data continuously monitored during the free-orbit phase of several space shuttle missions reveals global-scale distribution of ionospheric irregularities. Equatorial events cluster ±20 ◦ about the magnetic equator and polar events exhibit hemispheric asymmetry suggesting influence of an off-axis geomagnetic polar oval system. The diurnal, seasonal, and geographic distribution is compared to previous work concerning equatorial spread F, Appleton anomaly, and polar oval. The observations provide a succinct demonstration of the utility of space-based ionospheric monitoring using GPS. The susceptibility of GPS receivers to scintillation represents an unanticipated technical risk not factored into the selection of receivers for the U.S. space program.

The design and performance evaluation of an entry guidance algorithm for future space transportation vehicles is presented. The guidance concept is to plan and track aerodynamic acceleration. This concept, on which the longitudinal entry guidance for the Space Shuttle Orbiter is based, is extended to integrated longitudinal and lateral guidance. With integrated longitudinal and lateral guidance, more extreme points in the landing footprint can be reached accurately; in particular, the cross-range capability is extended. The guidance algorithm consists of two components: a trajectory planner and a trajectory tracking law. The planner generates reference drag acceleration and heading angle profiles, along with reference state and bank angle profiles. The planner executes onboard and is capable of generating updates as the entry evolves. The tracking law, based on feedback linearization, commands the angles of bank and attack required to follow the reference drag and heading angle profiles. The planner and tracking law are described, along with additional higher level logic included in the algorithm. Extensive simulations for a set of return-from-orbit entries, including ones requiring large cross range, demonstrate that this algorithm consistently achieves the desired target conditions within allowable tolerances and satisfies all other entry constraints.

Laboratory and on-orbit performance data for two common thermal control materials, silver- and aluminum-backed (metallized) fluorinated ethyl-propylene, were collected from a variety of sources and analyzed. It is demonstrated that the change in solar absorptance α is a strong function of particulate radiation for these materials. Examination of additional data shows that the atomic-oxygen recession rate is a strong function of solar exposure with an induction period of between 25 to 50 equivalent solar hours. The relationships determined in this analysis were incorporated into an electronic knowledge base, the "Spacecraft Materials Selector," under NASA Contract NAS8-98213. This tool is available from the NASA Space Environments and Effects program office.

In the choosing of a launch vehicle for a given mission or in the determination of insurance coverage and premiums fo rag iven launch, accurate estimates of the probability of success of the different launch vehicles provide important information. There are three general approaches to estimating the probability of launch success. The first is to use a probabilistic risk analysis, decomposing the system into its subsystems and components and estimating the probability of each of the failure modes. The second is to rely on expert judgment about the vehicle’s success rate as a whole, without a functional decomposition of the system. The third is to use statistical data about the past performance of the system to estimate the vehicle’s success rate. The focus is put on this last approach, using Bayesian probability theory to make better use of vehicle-level performance data. The procedure is demonstrated by an analysis of the success rates of most of the major families of launch vehicles currently in use in the world. A family of launch vehicles includes all variants of a particular type of vehicle from a specific manufacturer, for example, the Delta 2. For vehicles with a small number of launch attempts, the Bayesian approach provides the advantage over classic statistical approaches of yielding estimates of both the mean future frequency of success and the uncertainty about that mean.

Vibration reduction of space structures during launch is investigated experimentally. An optimal control law is formulated, based on a lumped parameter state-space model of the structure. A single piezoelectric actuator is bonded to the surface near the base to provide actuation. Band-limited white noise and recorded space shuttle launch excitations areused to assesstheperformance of thecontrol system. Signie cant reductionsin thevibrations are obtained using very low actuator power. The study indicates a large potential for improvements in payload protection during the space shuttle liftoff by using smart-structures technology. Nomenclature A = system matrix Aa = cross-sectional area of actuator, m 2 Ab = cross-sectional area of beam (wall), m 2

This paper presents results of flowfield calculations for typical hypersonic reentry conditions encountered by the nose region of the Space Shuttle Orbiter. Most of the transitional flow regime is covered by the altitude range of 150 to 92 km. Calculations were made with the Direct Simulation Monte Carlo (DSMC) method that accounts for translational, rotational, vibrational, and chemical nonequilibrium effects. Comparison of the DSMC heating results with both Shuttle flight data and continuum predictions showed good agreement at the lowest altitude considered. However, as the altitude increased, the continuum predictions, which did not include slip effects, departed rapidly from the DSMC results by overpredicting both heating and drag. The results demonstrate the effects of rarefaction on the shock and the shock layer, along with the extent of the slip and temperature jump at the surface. Also, the sensitivity of the flow structure to the gas-surface interaction model, thermal accommodation, and surface catalysis are studied.

The state of the art in aerodynamics engineering associated with the design and/or evaluation of reaction controls on flight vehicles operating in the atmosphere is reviewed. Various configurations of the aerodynamics interference problem are described to partition domains of the problem. To maintain an applications frame of reference, those descriptions are in the context of the dominant phenomenology observed under differing flight environment and vehicle geometry combinations. Following that, approaches to predicting or evaluating interference effects are reviewed. Approaches to the design of subscale wind-tunnel tests are discussed with a view toward relating appropriate scaling law approximations to the different domains of dominant phenomenology. Then results, found in the literature, describing the evolution of analytical and computational modeling are reviewed. Finally, some conclusions and observations on the state of the art are offered.

Real-time estimation of system states or parameters that are dife cult or expensive to measure directly is often needed for adaptive control or health monitoring purposes. A practical algorithm is proposed for adaptive state e ltering in nonlineardynamic systemswhen the state equations areunknown or too complex to model analytically. The state equationsare constructively approximated by using recurrent neural networks. The proposed algorithm is based on the predictor-update approach of the Kalman e lter, but a least-mean-square e lter implementation with an adaptive e lter gain is used. Furthermore, unlike the Kalman e lter and its nonlinear extensions, the proposed algorithm makes minimal assumptions regarding the underlying nonlinear system dynamics and their noise statistics. The e lter is used to estimate the high-pressure turbine discharge temperature of the space shuttle main engine, during setpoint changes and turbopump failures. The e lter is developed by using simulated engine data, and its performance is tested on both simulated and actual recorded space shuttle main engine transients. When the complexity of the problem studied is considered, the resulting e lter accuracy is shown to be quite acceptable. Further use of the adaptive e lter gain developed is to enable real-time detection of certain system failures, such as the turbopump failures of the space shuttle main engine. This concept is demonstrated also by using both simulated and experimental failure data.

As a result of the Wind spacecraft encounters with the moon, a new view of the lunar wake in the high-density solar wind plasma has emerged. Specifically, the lunar wake was considered to be magnetosonic in nature but is now demonstrated to be a kinetically driven structure filling in via ion sonic disturbances. The structure appears to be determined via kinetic plasma microinstabilities, rather than a bulk magnetohydrodynamic shock. Examining the specific structure, it becomes apparent that the lunar wake and that of the space shuttle have many similarities, suggesting that the shuttle wake is also driven via kinetic instabilities. Comparisons of the two wakes are presented in detail, illustrating the dominance of the kinetic phenomena in the replenishment of both plasma voids. The general concepts presented in this study have applications to other structures immersed in a plasma flow, including the space station.

Fully reusable two-stage-to-orbit launch vehicle designs that incorporate branching trajectories during their ascent are of current interest in the advanced launch vehicle design community. Unlike expendable vehicle designs, the booster of a two-stage reusable system must fly to a designated landing site after staging. Because of a mutual dependence on the staging conditions, both the booster flyback branch and the orbital branch of the ascent trajectory must be simultaneously optimized to achieve an overall system objective. The optimum solution is often a compromise between the local objectives of the two branches. Current and notable designs in this class include the U.S. Air Force Space Operations Vehicle designs, the Kelly Astroliner, the Kistler K-1, and NASA's proposed liquid flyback booster designs (space shuttle booster replacement). Solution techniques are introduced that are well suited to solving this class of problem with existing single-segment trajectory optimization codes. In particular, these methods originate from the field of multidisciplinary design optimization and include optimization-based decomposition and collaborative optimization. The results of applying these techniques to the branching trajectory optimization problem for the Kistler K-1 launch vehicle are given and conclusions are drawn with respect to computational efficiency and quality of the results. In general, partial optimization-based decomposition was preferred due to its superior robustness, ease of setup, fast execution time, and optimality of the results.

The orientation of a single-crystal material is known to affect the strength and life of structural component parts. Results are presented of an investigation of the effects of secondary axis orientation angles on the failure of the first-stage of the space shuttle main engine alternate turbopump development of the high-pressure fuel turbopump. First, the correlation of different failure models with low-cycle fatigue data for nickel-base single-crystal test specimens was analyzed. Then the models with the highest correlation coefficients were used to study the actual single-crystal blade structure. Based on the results obtained, a new failure model was proposed. A detailed finite element model for the first-stage blade was used to calculate the stresses and strains at all blade nodes for different material orientations. Results of the analysis showed that the critical value of the failure model could vary by up to a factor of 3 by changing the primary and secondary material orientations. A comparison between analytical results and engine test results showed good correlation and also demonstrated the dependence of cracking location on crystal orientation.

This research presents a new mathematical model characterizing the negative potential electrical charging behavior oflarge spacecraft inlowEarth orbitthatare actively collecting charge, such as partofanelectrodynamic tether system. The analysis was carried out to identify signie cant plasma current sources affecting steady-state spacecraft charging using data from the tethered satellite system missions. During both tethered satellite missions (Aug. 1992 and Feb. 1996), Space Shuttle Orbiter charging was lower than expected. The current collected by the Orbiter greatly exceeded premission predictions based on thin sheath, ram-dominated current collection. Our investigation revealed that the tethered satellite deployer boom was conducting and was grounded to the Orbiter providing asignie cantcurrent path from theplasma. Modelingresultssuggest thattheplasma sheath signie cantly augmented the ram current collected by the mainengine nozzles and the satellite deployer boom by expanding the effective current collecting area.

The new Space Shuttle superlightweight external fuel tank e ew for the e rst time on 2 June 1998 (Space Transportation System-mission 91 ). We present results of elastic linear-bifurcation buckling and nonlinear analyses of one of its major components; that is, the liquid-oxygen tank. The contents include an overview of the structure and a brief description of the e nite element code that was used to conduct the analyses. Results are presented that illustrate three distinctly different types of nonlinearresponsephenomena for thin-walled shells that aresubjected to combined mechanical and thermal loads that launch-vehicle shell designers may encounter. A procedure is demonstrated that can beused by structural analysts and designers to obtain reasonable, conservative estimates of linear-bifurcation, buckling-load knockdown factors for shells that are subjected to complex loading conditions or to characterize the effects of initial geometric imperfections on nonlinear shell response phenomena. Results are also presented that show that the superlightweight liquid-oxygen tank can carry loads in excess of twice the values of the operational prelaunch loads considered and that a e uid-e lled launch-vehicle shell can be highly sensitive to initial geometric imperfections. Presentedherearee vepapersonlarge-scaleanalysesofacomplexshellstructure.Thee rst,longerpaper,“ NonlinearBehaviorofSpaceShuttleSuperlightweight Liquid-OxygenTankUnderPrelaunchLoads,” coversthestructure,theanalysistechnique,someloading cases,andexperimentalverie cationonthemethod.Itis followed by a companion paper, “ Modeling and NonlinearStructural Analysis of a Large-ScaleLaunch Vehicle,” which goes into greaterdepth on themethods, and then by three shorter papers, “ Effects of Welding-Induced Imperfections on Behavior of Space Shuttle Superlightweight Tank,” “ Nonlinear Behavior of Space Shuttle Superlightweight Tank Under Booster Ascent Loads,” and “ Nonlinear Behavior of Space Shuttle Superlightweight Tank Under End-of-Flight Loads,” covering interesting behaviorunder differing load cases. Thesepapersare intended to stand ontheir own (and hencehavesomeredundant introductory material) but are complementary to one another, and so they are presented here together.

Results of linear-bifurcation and nonlinear analyses of the Space Shuttle superlightweight (SLWT) external liquid-oxygen (LO2) tank for an important early booster ascent loading condition are presented. These results for thin-walled linear elastic shells that are subjected to combined mechanical and thermal loads illustrate an important type of response mode that may be encountered in the design of other liquid-fuel launch vehicles. Linear-bifurcation analyses are presented that predict several nearly equal eigenvalues that correspond to local buckling modes in the forward ogive section of the LO 2 tank. In contrast, the nonlinear response phenomenon is shown to consist of short-wavelength bending deformations in the forward-ogive and barrel sections of the LO2 tank that grow in amplitude in a stable manner with increasing load. Imperfection sensitivity analyses are presented that show that the presence of several nearly equal eigenvalues does not lead to a premature general instability mode for the forward-ogive section. For the linear-bifurcation and nonlinear analyses, the results show that accurate predictions of the response of the shell generally require a large-scale, high-e delity, e nite element model, and that a design based on a linear-bifurcation buckling analysis and a buckling-load knockdown factor is overly conservative. Results are presented that show that the SLWT LO 2 tank can support loads in excess of approximately 2.6 timesthevaluesof theoperationalloads considered. In addition, results are presented that show that local bending deformations may cause failure of the thermal protection system (TPS) at load levels less than the load level corresponding to structural collapse. Results are presented that can be used to estimate the load level at which TPS failure is likely to occur.

Results of linear-bifurcation buckling and nonlinear analyses of the Space Shuttle superlightweight external liquid-oxygen tank are presented for an important prelaunch loading condition. These results show the effects of actual, measured welding-induced initial geometric imperfections on an important response mode for thin-walled shellsthataresubjectedtocombinedmechanicalandthermalloads.Thistypeofinitialgeometricimperfectionmay be encountered in the design of other liquid-fuel launch vehicles. Results are presented that show that the liquidoxygen tank will buckle in the barrel section, but at load levels nearly four times the magnitude of the operational load level, and will exhibit stable postbuckling behavior. The actual measured imperfections are located in this sectionofthetank. Resultsofimperfection sensitivity analysesarepresented thatshowthat thelargestdegradation in the apparent membrane stiffnesses of the liquid-oxygen tank barrel section is caused by an imperfection shape that is in the form of the linear-bifurcation buckling mode with a relatively small amplitude. These results also show that the effect of the relatively large-amplitude measured imperfection is benign.