ALBERT

Beschreibung: Maintaining contamination certification of multi-mission flight hardware is an innovative approach to controlling mission costs. Methods for assessing ground induced degradation between missions have been employed by the Hubble Space Telescope (HST) Project for the multi-mission (servicing) hardware. By maintaining the cleanliness of the hardware between missions, and by controlling the materials added to the hardware during modification and refurbishment both project funding for contamination recertification and schedule have been significantly reduced. These methods will be discussed and HST hardware data will be presented.

Beschreibung: The Earth Observing System (EOS) AM-1 spacecraft for NASA's Mission to Planet Earth is scheduled to be launched on an Atlas IIAS vehicle in June of 1998. One concern is that the instruments on the EOS spacecraft are sensitive to the shock-induced vibration produced when the spacecraft separates from the launch vehicle. By employing unique statistical analysis to the available ground test shock data, the NASA Lewis Research Center found that shock-induced vibrations would not be as great as the previously specified levels of Lockheed Martin. The EOS pyroshock separation testing, which was completed in 1997, produced a large quantity of accelerometer data to characterize the shock response levels at the launch vehicle/spacecraft interface. Thirteen pyroshock separation firings of the EOS and payload adapter configuration yielded 78 total measurements at the interface. The multiple firings were necessary to qualify the newly developed Lockheed Martin six-hardpoint separation system. Because of the unusually large amount of data acquired, Lewis developed a statistical methodology to predict the maximum expected shock levels at the interface between the EOS spacecraft and the launch vehicle. Then, this methodology, which is based on six shear plate accelerometer measurements per test firing at the spacecraft/launch vehicle interface, was used to determine the shock endurance specification for EOS. Each pyroshock separation test of the EOS spacecraft simulator produced its own set of interface accelerometer data. Probability distributions, histograms, the median, and higher order moments (skew and kurtosis) were analyzed. The data were found to be lognormally distributed, which is consistent with NASA pyroshock standards. Each set of lognormally transformed test data produced was analyzed to determine if the data should be combined statistically. Statistical testing of the data's standard deviations and means (F and t testing, respectively) determined if data sets were significantly different at a 95-percent confidence level. If two data sets were found to be significantly different, these families of data were not combined for statistical purposes. This methodology produced three separate statistical data families of shear plate data. For each population, a P99.1/50 (probability/confidence) per-separation-nut firing level was calculated. By using the binomial distribution, Lewis researchers determined that this pernut firing level was equivalent to a P95/50 per-flight confidence level. The overall envelope of the per-flight P95/50 levels led to Lewis' recommended EOS interface shock endurance specification. A similar methodology was used to develop Lewis' recommended EOS mission assurance levels. The available test data for the EOS mission are significantly larger than for a normal mission, thus increasing the confidence level in the calculated expected shock environment. Lewis significantly affected the EOS mission by properly employing statistical analysis to the data. This analysis prevented a costly requalification of the spacecraft's instruments, which otherwise would have been exposed to significantly higher test levels.

Beschreibung: NASA Lewis Research Center recently led a multi-organizational acoustic test program. This testing consisted of acoustically exciting a Cassini spacecraft simulator in a full scale 60 foot high Titan 4 payload fairing with various acoustic blanket designs and configurations in a large reverberant acoustic chamber. The primary purpose of this test program was to measure the fairing's internal acoustics and spacecraft vibration, especially the Radioisotope Thermoelectric Generators (RTG) vibration, and to quantify the mitigation efforts in reducing these levels. Key to this reduction effort was the utilization of new acoustic blanket designs. This paper will provide the background and rationale for performing this test program, state the test program's primary and secondary objectives and describe the test matrix, hardware and instrumentation. A second part companion paper will provide the test results and data analysis.

Beschreibung: A Cassini spacecraft simulator in a full scale 60 foot high Titan 4 payload fairing with various acoustic blanket designs and configurations was recently tested in a large reverberant acoustic chamber. A first part companion paper provides the test configuration details and other background information. This paper addresses the results obtained from this test program. Emphasis will be on the effects of the new blanket designs on reducing the payload fairing's internal acoustics and the vibration response of the spacecraft's Radioisotope Thermoelectric Generators. Other results discussed include: the effect of blankets on fairing vibration, the effect of partial blanket coverage on acoustics and vibration and the effect of tuned vibration absorbers.

Beschreibung: One of the most interesting and challenging aspects of formation guidance law design is the coupling of the orbit design and the science return. The analyst s role is more complicated than simply to design the formation geometry and evolution. He or she is also involved in designing a significant portion of the science instrument itself. The effectiveness of the formation as a science instrument is intimately coupled with the relative geoniet,ry and evolution of the collection of spacecraft. Therefore, the science return can be maximized by optimizing the orbit design according to a performance metric relevant to the science mission goals. In this work, we present a simple method for optimal formation guidance that is applicable to missions whose performance metric, requirements, and constraints can be cast as functions that are explicitly dependent upon the orbit states and spacecraft relative positions and velocities. We present a general form for the cost and constraint functions, and derive their semi-analytic gradients with respect to the formation initial conditions. The gradients are broken down into two types. The first type are gradients of the mission specific performance metric with respect to formation geometry. The second type are derivatives of the formation geometry with respect to the orbit initial conditions. The fact that these two types of derivatives appear separately allows us to derive and implement a general framework that requires minimal modification to be applied to different missions or mission phases. To illustrate the applicability of the approach, we conclude with applications to twc missims: the Magnetospheric Mu!tiscale mission (MMS), a,nd the TJaser Interferometer Space Antenna (LISA).

Beschreibung: The Inflatable Re-entry Vehicle Experiment (IRVE) is a 3.0 meter, 60 degree half-angle sphere cone, inflatable aeroshell experiment designed to demonstrate various aspects of inflatable technology during Earth re-entry. IRVE will be launched on a Terrier-Improved Orion sounding rocket from NASA s Wallops Flight Facility in the fall of 2006 to an altitude of approximately 164 kilometers and re-enter the Earth s atmosphere. The experiment will demonstrate exo-atmospheric inflation, inflatable structure leak performance throughout the flight regime, structural integrity under aerodynamic pressure and associated deceleration loads, thermal protection system performance, and aerodynamic stability. Structural integrity and dynamic response of the inflatable will be monitored with photogrammetric measurements of the leeward side of the aeroshell during flight. Aerodynamic stability and drag performance will be verified with on-board inertial measurements and radar tracking from multiple ground radar stations. In addition to demonstrating inflatable technology, IRVE will help validate structural, aerothermal, and trajectory modeling and analysis techniques for the inflatable aeroshell system. This paper discusses the structural analysis and testing of the IRVE inflatable structure. Equations are presented for calculating fabric loads in sphere cone aeroshells, and finite element results are presented which validate the equations. Fabric material properties and testing are discussed along with aeroshell fabrication techniques. Stiffness and dynamics tests conducted on a small-scale development unit and a full-scale prototype unit are presented along with correlated finite element models to predict the in-flight fundamental mod

Beschreibung: Inflatable aeroshells offer several advantages over traditional rigid aeroshells for atmospheric entry. Inflatables offer increased payload volume fraction of the launch vehicle shroud and the possibility to deliver more payload mass to the surface for equivalent trajectory constraints. An inflatable s diameter is not constrained by the launch vehicle shroud. The resultant larger drag area can provide deceleration equivalent to a rigid system at higher atmospheric altitudes, thus offering access to higher landing sites. When stowed for launch and cruise, inflatable aeroshells allow access to the payload after the vehicle is integrated for launch and offer direct access to vehicle structure for structural attachment with the launch vehicle. They also offer an opportunity to eliminate system duplication between the cruise stage and entry vehicle. There are however several potential technical challenges for inflatable aeroshells. First and foremost is the fact that they are flexible structures. That flexibility could lead to unpredictable drag performance or an aerostructural dynamic instability. In addition, durability of large inflatable structures may limit their application. They are susceptible to puncture, a potentially catastrophic insult, from many possible sources. Finally, aerothermal heating during planetary entry poses a significant challenge to a thin membrane. NASA Langley Research Center and NASA's Wallops Flight Facility are jointly developing inflatable aeroshell technology for use on future NASA missions. The technology will be demonstrated in the Inflatable Re-entry Vehicle Experiment (IRVE). This paper will detail the development of the initial IRVE inflatable system to be launched on a Terrier/Orion sounding rocket in the fourth quarter of CY2005. The experiment will demonstrate achievable packaging efficiency of the inflatable aeroshell for launch, inflation, leak performance of the inflatable system throughout the flight regime, structural integrity when exposed to a relevant dynamic pressure and aerodynamic stability of the inflatable system. Structural integrity and structural response of the inflatable will be verified with photogrammetric measurements of the back side of the aeroshell in flight. Aerodynamic stability as well as drag performance will be verified with on board inertial measurements and radar tracking from multiple ground radar stations. The experiment will yield valuable information about zero-g vacuum deployment dynamics of the flexible inflatable structure with both inertial and photographic measurements. In addition to demonstrating inflatable technology, IRVE will validate structural, aerothermal, and trajectory modeling techniques for the inflatable. Structural response determined from photogrammetrics will validate structural models, skin temperature measurements and additional in-depth temperature measurements will validate material thermal performance models, and on board inertial measurements along with radar tracking from multiple ground radar stations will validate trajectory simulation models.

Beschreibung: Maintaining contamination certification of multi-mission flight hardware is an innovative approach to controlling mission costs. Methods for assessing ground induced degradation between missions have been employed by the Hubble Space Telescope (HST) Project for the multi-mission (servicing) hardware. By maintaining the cleanliness of the hardware between missions, and by controlling the materials added to the hardware during modification and refurbishment both project funding for contamination recertification and schedule have been significantly reduced. These methods will be discussed and HST hardware data will be presented.

Beschreibung: The Ares I launch vehicle will be NASA s first new launch vehicle since 1981. Currently in design, it will replace the Space Shuttle in taking astronauts to the International Space Station, and will eventually play a major role in humankind s return to the Moon and eventually to Mars. Prior to any manned flight of this vehicle, unmanned test readiness flights will be flown. The first of these readiness flights, named Ares I-X, is scheduled to be launched in April 2009. The NASA Glenn Research Center is responsible for the design, manufacture, test and analysis of the Ares I-X upper stage simulator (USS) element. As part of the design effort, the structural dynamic response of the Ares I-X launch vehicle to its vibroacoustic flight environments must be analyzed. The launch vehicle will be exposed to extremely high acoustic pressures during its lift-off and aerodynamic stages of flight. This in turn will cause high levels of random vibration on the vehicle's outer surface that will be transmitted to its interior. Critical flight equipment, such as its avionics and flight guidance components are susceptible to damage from this excitation. This study addresses the modelling, analysis and predictions from examining the structural dynamic response of the Ares I-X upper stage to its vibroacoustic excitations. A statistical energy analysis (SEA) model was used to predict the high frequency response of the vehicle at locations of interest. Key to this study was the definition of the excitation fields corresponding to lift off acoustics and the unsteady aerodynamic pressure fluctuations during flight. The predicted results will be used by the Ares I-X Project to verify the flight qualification status of the Ares I-X upper stage components.

Beschreibung: This course will provide students with the understanding necessary to calculate certification requirements for and interpret the data from hardware undergoing vacuum bakeout. The fundamental physics that driv e hardware outgassing and contaminant deposition will be explained, a nd data reduction techniques will be applied to examples.