SCIENTIFIC AND TECHNICAL AEROSPACE REPORTS
A Biweekly Publication of the National Aeronautics and Space Administration
VOLUME 44, ISSUE 9 - MAY 5, 2006
18 SPACECRAFT DESIGN, TESTING AND PERFORMANCE
Includes satellites; space platforms; space stations; spacecraft systems and components such as thermal and environmental controls; and spacecraft control and stability characteristics.
For life support systems see 54 Man/System Technology and Life Support.
For related information see also 05 Aircraft Design, Testing and Performance; 39 Structural Mechanics; and 16 Space Transportation and Safety.
20060012088 NASA Langley Research Center, Hampton, VA, USA
Aerocapture Systems Analysis for a Neptune Mission
Lockwood, Mary Kae; Edquist, Karl T.; Starr, Brett R.; Hollis, Brian R.; Hrinda, Glenn A.; Bailey, Robert W.; Hall, JefferyL.; Spilker, Thomas R.; Noca, Muriel A.; O’Kongo, N., et al.; April 2006; 148 pp.; In English; See also 20060012089 -20060012100; Original contains color and black and white illustrationsContract(s)/Grant(s): 620.02.01.01.07.02Report No.(s): NASA/TM-2006-214300; L-19236; Copyright; Avail.: CASI: A07, Hardcopy
A Systems Analysis was completed to determine the feasibility, benefit and risk of an aeroshell aerocapture system forNeptune and to identify technology gaps and technology performance goals. The systems analysis includes the followingdisciplines: science; mission design; aeroshell configuration; interplanetary navigation analyses; atmosphere modeling;computational fluid dynamics for aerodynamic performance and aeroheating environment; stability analyses; guidancedevelopment; atmospheric flight simulation; thermal protection system design; mass properties; structures; spacecraft designand packaging; and mass sensitivities. Results show that aerocapture is feasible and performance is adequate for the Neptunemission. Aerocapture can deliver 1.4 times more mass to Neptune orbit than an all-propulsive system for the same launchvehicle and results in a 3-4 year reduction in trip time compared to all-propulsive systems. Enabling technologies for thismission include TPS manufacturing; and aerothermodynamic methods for determining coupled 3-D convection, radiation andablation aeroheating rates and loads.Author
Systems Analysis; Aeroshells; Aerocapture; Atmospheric Models; Computational Fluid Dynamics; Neptune (Planet);Spacecraft Design; Systems Engineering
20060012089 NASA Johnson Space Center, Houston, TX, USA
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Aerocapture Guidance Performance for the Neptune Orbiter
Masciarelli, James P.; Westhelle, Carlos H.; Graves, Claude A.; Aerocapture Systems Analysis for a Neptune Mission; April 2006, pp. 98-106; In English; See also 20060012088; Original contains color and black and white illustrations; Copyright; Avail.: CASI: A02, Hardcopy
A performance evaluation of the Hybrid Predictor-corrector Aerocapture Scheme (HYPAS) guidance algorithm for aerocapture at Neptune is presented in this paper for a Mission to Neptune and the Neptune moon Triton. This mission has several challenges not experienced in previous aerocapture guidance assessments. These challengers are a very high Neptune arrival speed, atmospheric exit into a high energy orbit about Neptune, and a very high ballistic coefficient that results in a low altitude acceleration capability when combined with the aeroshell L/D. The evaluation includes a definition of the entry corridor, a comparison to the theoretical optimum performance, and guidance responses to variations in atmospheric density, aerodynamic coefficients and flight path angle for various vehicle configurations (ballistic numbers). The benefits of utilizing angle-of-attack modulation in addition to bank angle modulation to improve flight performance is also discussed. The results show that despite large sensitivities in apoapsis targeting, the algorithm performs within the allocated (Delta)V budget for the Neptune mission using only bank angle modulation. The addition of angle-of-attack modulation with as little as +/-5 degrees of amplitude significantly improves the accuracy in final orbit apoapsis. Although angle-of-attack modulation complicates the vehicle design its performance enhancement reduces aerocapture risk and reduces the propellant consumption needed to reach the high energy target orbit. Author
Flight Characteristics; Performance Tests; Predictor-Corrector Methods; Flight Paths; Aerodynamic Coefficients; Aerocapture; Aeroshells; Angle of Attack; Atmospheric Density
20060012090 NASA Langley Research Center, Hampton, VA, USA
Preliminary Convective-Radiative Heating Environments for a Neptune Aerocapture Mission
Hollis, Brian R.; Wright, Michael J.; Olejniczak, Joseph; Takashima, Naruhisa; Sutton, Kenneth; Prabhu, Dinesh; Aerocapture Systems Analysis for a Neptune Mission; April 2006, pp. 107-118; In English; See also 20060012088; Original contains color illustrations Contract(s)/Grant(s): NAS2-99092; NAS1-00135; NCC1-02043; Copyright; Avail.: CASI: A03, Hardcopy
Convective and radiative heating environments have been computed for a three-dimensional ellipsled configuration which would perform an aerocapture maneuver at Neptune. This work was performed as part of a one-year Neptune aerocapture spacecraft systems study that also included analyses of trajectories, atmospheric modeling, aerodynamics, structural design, and other disciplines. Complementary heating analyses were conducted by separate teams using independent sets of aerothermodynamic modeling tools (i.e. Navier-Stokes and radiation transport codes). Environments were generated for a large 5.50 m length ellipsled and a small 2.88 m length ellipsled. Radiative heating was found to contribute up to 80% of the total heating rate at the ellipsled nose depending on the trajectory point. Good agreement between convective heating predictions from the two Navier-Stokes solvers was obtained. However, the radiation analysis revealed several uncertainties in the computational models employed in both sets of codes, as well as large differences between the predicted radiative heating rates. Author
Aerocapture; Convective Heat Transfer; Navier-Stokes Equation; Radiation Transport; Aerothermodynamics; Aerodynamics; Atmospheric Models
20060012091 Jet Propulsion Lab., California Inst. of Tech., Pasadena, CA, USA
Aerocapture Navigation at Neptune
Haw, Robert J.; Aerocapture Systems Analysis for a Neptune Mission; April 2006, pp. 57-73; In English; See also 20060012088; Original contains color and black and white illustrations; Copyright; Avail.: CASI: A03, Hardcopy
A proposed Neptune Orbiter Aerocapture mission will use solar electric propulsion to send an orbiter to Neptune. Navigation feasibility of direct-entry aerocapture for orbit insertion at Neptune is shown. The navigation strategy baselines optical imaging and VLBI measurements in order to satisfy the flight system s atmosphere entry flight path angle, which is targeted to enter Neptune with an entry flight path angle of 11.6 degrees. Error bars on the entry flight path angle of plus or minus 0.55 (3 ) are proposed. This requirement can be satisfied with a data cutoff 3.2 days prior to arrival. There is some margin in the arrival template to tighten (i.e. reduce) the entry corridor either by scheduling a data cutoff closer to Neptune or alternatively, reducing uncertainties by increasing the fidelity of the optical navigation camera. Author
Aerocapture; Navigation; Neptune (Planet); Space Missions; NASA Space Programs
20060012092 NASA Langley Research Center, Hampton, VA, USA
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Neptune Aerocapture Systems Analysis
Lockwood, Mary Kae; Aerocapture Systems Analysis for a Neptune Mission; April 2006, pp. 1-16; In English; See also 20060012088; Original contains color and black and white illustrations; Copyright; Avail.: CASI: A03, Hardcopy
A Neptune Aerocapture Systems Analysis is completed to determine the feasibility, benefit and risk of an aeroshell aerocapture system for Neptune and to identify technology gaps and technology performance goals. The high fidelity systems analysis is completed by a five center NASA team and includes the following disciplines and analyses: science; mission design; aeroshell configuration screening and definition; interplanetary navigation analyses; atmosphere modeling; computational fluid dynamics for aerodynamic performance and database definition; initial stability analyses; guidance development; atmospheric flight simulation; computational fluid dynamics and radiation analyses for aeroheating environment definition; thermal protection system design, concepts and sizing; mass properties; structures; spacecraft design and packaging; and mass sensitivities. Results show that aerocapture can deliver 1.4 times more mass to Neptune orbit than an all-propulsive system for the same launch vehicle. In addition aerocapture results in a 3-4 year reduction in trip time compared to all-propulsive systems. Aerocapture is feasible and performance is adequate for the Neptune aerocapture mission. Monte Carlo simulation results show 100% successful capture for all cases including conservative assumptions on atmosphere and navigation. Enabling technologies for this mission include TPS manufacturing; and aerothermodynamic methods and validation for determining coupled 3-D convection, radiation and ablation aeroheating rates and loads, and the effects on surface recession. Author
Aerocapture; Aerothermodynamics; Computational Fluid Dynamics; Interplanetary Navigation; Navigation; Mission Planning; Neptune (Planet); Aeroshells
20060012093 Jet Propulsion Lab., California Inst. of Tech., Pasadena, CA, USA
Neptune Aerocapture Mission and Spacecraft Design Overview
Bailey, R. W.; Hall, J. L.; Spilker, T. R.; O’Kongo, N.; Aerocapture Systems Analysis for a Neptune Mission; April 2006, pp. 17-28; In English; See also 20060012088; Original contains color and black and white illustrations; Copyright; Avail.: CASI: A03, Hardcopy
A detailed Neptune aerocapture systems analysis and spacecraft design study was performed as part of NASA s In-Space Propulsion Program. The primary objective was to assess the feasibility of a spacecraft point design for a Neptune/Triton science mission that uses aerocapture as the Neptune orbit insertion mechanism. This paper provides an overview of the science, mission and spacecraft design resulting from that study. The estimated delivered wet mass allocation to Neptune orbit was approximately 928 kg. The aerocapture entry system, comprised of aeroshell and post-aerocapture orbit correction propellant, was approximately 1252 kg, for a total atmospheric entry mass allocation of approximately 2239 kg. The aeroshell used was a 2.88 m long flattened ellipsled with a lift to drag ratio of 0.8. A Delta-IV Heavy launch vehicle combined with a 30kW solar electric propulsion (SEP) stage and a Venus/Jupiter gravity assist were used to get the spacecraft to Neptune in 10.25 years. The SEP stage and Orbiter both have 35% dry mass margins ((allocation CBE)/allocation) and the overall launch stack has an additional approximately 8% unallocated reserve. The feasibility of the mission requires the solution of two key technical challenges: improvement in aerothermodynamic computational tools for Neptune; and development of thermal protection material manufacturing processes for the increased thickness needed for aerocapture. Several other component technologies were identified as being able to provide significant performance improvements including: radioiostopic power generation, solar cells and array structure, low mass/power science instruments, and small stowed volume/large aperture deployable Ka-Band antennas. Author
Aerocapture; Neptune (Planet); Spacecraft Design; General Overviews; Space Missions; Systems Analysis
20060012094 Jet Propulsion Lab., California Inst. of Tech., Pasadena, CA, USA
Mission Trades for Aerocapture at Neptune
Noca, Muriel A.; Bailey, RobertW.; Aerocapture Systems Analysis for a Neptune Mission; April 2006, pp. 29-44; In English; See also 20060012088; Original contains color and black and white illustrations; Copyright; Avail.: CASI: A03, Hardcopy
A detailed Neptune aerocapture systems analysis and spacecraft design study was performed to improve our understanding of the technology requirement for such a hard mission. The primary objective was to engineer a point design based on blunt body aeroshell technology and quantitatively assess feasibility and performance. This paper reviews the launch vehicle, propulsion, and trajectory options to reach Neptune in the 2015-2020 time frame using aerocapture and all-propulsive vehicles. It establishes the range of entry conditions that would be consistent with delivering a approx. 1900 kg total entry vehicle maximum expected mass to Neptune including a approx. 790 kg orbiter maximum expected mass to the science orbit. Two Neptune probes would be also be delivered prior to the aerocapture maneuver. Results show that inertial entry velocities in the range of 28 to 30 km/s are to be expected for chemical and solar electric propulsion options with several gravity assists (combinations of Venus, Earth and Jupiter gravity assists). Trip times range from approximately 10-11 years for aerocapture orbiters to 15 years for all-propulsive vehicles. This paper shows that the use of aerocapture enables this mission given the payload to deliver around Neptune compared to an all-propulsive orbit insertion approach. However, an all-propulsive chemical insertion option is possible for lower payload masses than the one needed for this science mission. Both approaches require a Delta IV heavy class launch vehicle. Author
Aerocapture; Aeroshells; Systems Analysis; Spacecraft Design; Blunt Bodies; Launch Vehicles; Propulsion; Trajectories
20060012095 NASA Langley Research Center, Hampton, VA, USA
Configuration, Aerodynamics, and Stability Analysis for a Neptune Aerocapture Orbiter
Edquist, Karl T.; Prabhu, Ramadas K.; Hoffman, David A.; Rea, Jeremy R.; Aerocapture Systems Analysis for a Neptune Mission; April 2006, pp. 45-56; In English; See also 20060012088; Original contains color illustrations; Copyright; Avail.: CASI: A03, Hardcopy
A multi-center NASA team conducted a systems analysis study of a Neptune aerocapture orbiter mission in order to demonstrate feasibility and identify technology gaps. The aerocapture maneuver utilizes aerodynamic drag to decelerate the vehicle, rather than chemical propulsion, for orbit insertion around Neptune and allows a flyby of the Triton moon. This paper presents the analysis used to select an orbiter shape, and the aerodynamics and stability characteristics of the reference vehicle. Several shape classes were screened for aerodynamic performance using modified Newtonian theory. A lift-to-drag ratio requirement of 0.6 to 0.8 was derived from an estimate of the theoretical corridor width to give margin beyond 3-sigma dispersions. A flat-bottomed ellipsled was selected as the reference orbiter shape based on various metrics, including lift-to-drag ratio, ballistic coefficient, and effective volume. High-fidelity computational solutions for the reference orbiter shape predict a lift-to-drag ratio of 0.806 and ballistic coefficient of 895 kg/sq m at a trim angle-of-attack of 40 deg. Stable pitch behavior is predicted with a 6.2% static margin for an axial center of gravity at 51% of the vehicle length from the nose. Both the longitudinal short-period and lateral Dutch-roll frequencies are shown to be within acceptable limits based on piloted vehicle specifications. Aerodynamics uncertainties were estimated to result in a lift-to-drag ratio uncertainty of +13.4%/- 14.5% using RSS values and +26.1%/-22.2% using stacked worst-case values. Author
Aerocapture; Feasibility; Systems Analysis; Flyby Missions; Aerodynamic Characteristics; Aerodynamic Configurations; Aerodynamic Stability; Angle of Attack
20060012096 NASA Marshall Space Flight Center, Huntsville, AL, USA
Atmospheric Models for Aerocapture
Justus, C. G.; Duvall, Aleta L.; Keller, Vernon W.; Aerocapture Systems Analysis for a Neptune Mission; April 2006, pp. 74-79; In English; See also 20060012088; Original contains color illustrations; Copyright; Avail.: CASI: A02, Hardcopy
There are eight destinations in the Solar System with sufficient atmosphere for aerocapture to be a viable aeroassist option: Venus, Earth, Mars, Jupiter, Saturn and its moon Titan, Uranus, and Neptune. Engineering-level atmospheric models for four of these targets - Earth, Mars, Titan, and Neptune - have been developed for NASA to support systems analysis studies of potential future aerocapture missions. Development of a similar atmospheric model for Venus has recently commenced. An important capability of all of these models is their ability to simulate quasi-random density perturbations for Monte Carlo analysis in developing guidance, navigation and control algorithms, and for thermal systems design. Similarities and differences among these atmospheric models are presented, with emphasis on the recently developed Neptune model and on planned characteristics of the Venus model. Example applications for aerocapture are also presented and illustrated. Recent updates to the Titan atmospheric model are discussed, in anticipation of application to trajectory and atmospheric reconstruct for the Huygens Probe entry at Titan. Author
Aerocapture; Atmospheric Models; Solar System; Systems Analysis
20060012097 NASA Marshall Space Flight Center, Huntsville, AL, USA
Atmospheric Models for Aerocapture Systems Studies
Justus, C. G.; Duvall, Aleta L.; Keller, Vernon W.; Aerocapture Systems Analysis for a Neptune Mission; April 2006, pp. 80-86; In English; See also 20060012088; Original contains color illustrations; Copyright; Avail.: CASI: A02, Hardcopy
Aerocapture uses atmospheric drag to decelerate into captured orbit from interplanetary transfer orbit. This includes capture into Earth orbit from, for example, Lunar-return or Mars-return orbit. Eight Solar System destinations have sufficient atmosphere for aerocapture to be applicable three of the rocky planets (Venus, Earth, and Mars), four gas giants (Jupiter, Saturn, Uranus, and Neptune), and Saturn s moon Titan. These destinations fall into two general groups: (1) The rocky planets, which have warm surface temperatures (about 200 K to 750 K) and rapid decrease of density with altitude, and (2) the gas giants and Titan, which have cold temperatures (about 70 K to 170 K) at the surface or 1-bar pressure level, and slow rate of decrease of density with altitude.
Aerocapture altitudes at the gas giants typically range from about 150 km to 300 km above the 1-bar pressure reference. Aerocapture at the rocky planets would occur at altitudes of about 50 km to 100 km. In contrast, aerobraking (circularizing a highly elliptical capture orbit, using multiple atmospheric passes) would occur at widely varying altitudes ranging from about 125 km out to 700 km for Titan. In addition to aerocapture altitude, aerocapture corridor width is also determined by details of the atmospheric density profile. Corridor width is the range of atmospheric entry angles allowable for successful aerocapture, i.e., achieving capture orbit without skip-out or burn-up . Corridor width is significantly affected by rate of change of density, as measured by the density scale height, at aerocapture periapsis altitude. Density scale height is the vertical distance over which density changes by a factor of e, (i.e., logarithmically). Larger scale height values mean slower density variation with height and larger corridor width; smaller scale height leads to smaller corridor width.
For the rocky planets, the overall rapid fall-off of density with height leads to relatively low density scale heights at aerocapture altitudes and small aerocapture corridor widths for these destinations. Larger density scale heights, with consequently larger corridor widths, result from the slower density fall-of with height for Titan and the gas giant planets. Density scale height values at periapsis for the rocky planets vary from about 4 km to 8 km; for the gas giant planets and Titan this range is about 25 km to 50 km. Engineering-level atmospheric models for Earth, Mars, Titan, and Neptune have been developed for NASA systems analysis studies of potential future aerocapture missions. Development of a similar atmospheric model for Venus has recently commenced.
These models are collectively referred to as Global Reference Atmosphere Models, or GRAMs. An important capability of all of the GRAM models is their ability to simulate quasi-random density perturbations for Monte Carlo analyses in developing guidance, navigation, and control algorithms, and for thermal systems design. Small-scale root-mean-square (rms) density perturbations observed for Earth may be compared with those modeled for Mars, Titan, and Neptune. Monte-Carlo simulations of density variations for Neptune atmospheric conditions yield minimum, average, and maximum density profiles due to expected variations with season, latitude, time-of-day, etc. Details of these comparisons and simulations are discussed. Author
Aerocapture; Atmospheric Density; Interplanetary Transfer Orbits; Atmospheric Entry; Systems Analysis; Surface Temperature; Elliptical Orbits; Corridors; Atmospheric Models; Aerobraking
20060012098 NASA Langley Research Center, Hampton, VA, USA
Aerocapture Performance Analysis for a Neptune-Triton Exploration Mission
Starr, Brett R.; Westhelle, Carlos H.; Masciarelli, James P.; Aerocapture Systems Analysis for a Neptune Mission; April 2006, pp. 87-97; In English; See also 20060012088; Original contains color and black and white illustrations; Copyright; Avail.: CASI: A03, Hardcopy
A systems analysis has been conducted for a Neptune-Triton Exploration Mission in which aerocapture is used to capture a spacecraft at Neptune. Aerocapture uses aerodynamic drag instead of propulsion to decelerate from the interplanetary approach trajectory to a captured orbit during a single pass through the atmosphere. After capture, propulsion is used to move the spacecraft from the initial captured orbit to the desired science orbit. A preliminary assessment identified that a spacecraft with a lift to drag ratio of 0.8 was required for aerocapture. Performance analyses of the 0.8 L/D vehicle were performed using a high fidelity flight simulation within a Monte Carlo executive to determine mission success statistics. The simulation was the Program to Optimize Simulated Trajectories (POST) modified to include Neptune specific atmospheric and planet models, spacecraft aerodynamic characteristics, and interplanetary trajectory models. To these were added autonomous guidance and pseudo flight controller models. The Monte Carlo analyses incorporated approach trajectory delivery errors, aerodynamic characteristics uncertainties, and atmospheric density variations. Monte Carlo analyses were performed for a reference set of uncertainties and sets of uncertainties modified to produce increased and reduced atmospheric variability. For the reference uncertainties, the 0.8 L/D flatbottom ellipsled vehicle achieves 100% successful capture and has a 99.87 probability of attaining the science orbit with a 360 m/s DelatV budget for apoapsis and periapsis adjustment. Monte Carlo analyses were also performed for a guidance system that modulates both bank angle and angle of attack with the reference set of uncertainties. An alpha and bank modulation guidance system reduces the 99.87 percentile DeltaV 173 m/s (48%) to 187 m/s for the reference set of uncertainties. Author
Aerocapture; Neptune (Planet); Triton; Space Exploration; Systems Analysis; Space Missions
20060012099 NASA Ames Research Center, Moffett Field, CA, USA
TPS Challenges for Neptune Aerocapture
Laub, B.; Chen, Y. K.; Aerocapture Systems Analysis for a Neptune Mission; April 2006, pp. 119-129; In English; See also 20060012088; Original contains color and black and white illustrations; Copyright; Avail.: CASI: A03, Hardcopy
A study to develop a conceptual design for an aerocapture mission at Neptune was conducted by a NASA systems analysis team comprised of technical experts from several NASA centers. Multidisciplinary analyses demonstrated that aerocapture could be accomplished at Neptune with a rigid aeroshell with a flattened ellipsled geometry flying at a nominal angle-of-attack of 40 degrees entering the Neptune atmosphere at an inertial entry velocity of approx. equals 29 km/s. Aerothermal analyses demonstrated that both the peak convective and radiative heating rates in the stagnation region are very severe. Furthermore, due to the duration of the aerocapture trajectory, the total integrated heat loads are extremely large. TPS sizing analyses were conducted for a limited range of candidate TPS materials since such high peak heat fluxes limit candidate materials to dense, carbonaceous ablators. On the windward side, in regions away from the stagnation region, lower density ablators may suffice. Low-density ablators also are viable candidates on the lee side. However, there are significant uncertainties associated with the turbulent and radiative heating rates. TPS mass requirements for this mission are very large, and the ablator thickness requirements may be beyond current manufacturing capabilities. Author
Aerocapture; Systems Analysis; Aeroshells; Atmospheric Entry; Angle of Attack; Neptune Atmosphere; Neptune (Planet); Stagnation Point; Convective Heat Transfer
20060012100 NASA Langley Research Center, Hampton, VA, USA
Structural Design for a Neptune Aerocapture Mission
Dyke, R. Eric; Hrinda, Glenn A.; Aerocapture Systems Analysis for a Neptune Mission; April 2006, pp. 130-142; In English; See also 20060012088; Original contains color and black and white illustrations; Copyright; Avail.: CASI: A03, Hardcopy
A multi-center study was conducted in 2003 to assess the feasibility of and technology requirements for using aerocapture to insert a scientific platform into orbit around Neptune. The aerocapture technique offers a potential method of greatly reducing orbiter mass and thus total spacecraft launch mass by minimizing the required propulsion system mass. This study involved the collaborative efforts of personnel from Langley Research Center (LaRC), Johnson Space Flight Center (JSFC), Marshall Space Flight Center (MSFC), Ames Research Center (ARC), and the Jet Propulsion Laboratory (JPL). One aspect of this effort was the structural design of the full spacecraft configuration, including the ellipsled aerocapture orbiter and the in-space solar electric propulsion (SEP) module/cruise stage. This paper will discuss the functional and structural requirements for each of these components, some of the design trades leading to the final configuration, the loading environments, and the analysis methods used to ensure structural integrity. It will also highlight the design and structural challenges faced while trying to integrate all the mission requirements. Component sizes, materials, construction methods and analytical results, including masses and natural frequencies, will be presented, showing the feasibility of the resulting design for use in a Neptune aerocapture mission. Lastly, results of a post-study structural mass optimization effort on the ellipsled will be discussed, showing potential mass savings and their influence on structural strength and stiffness Author
Structural Design; Spacecraft Configurations; Aerocapture; Functional Design Specifications; Solar Electric Propulsion
20060012130 NASA Johnson Space Center, Houston, TX, USA
Estimation of the Space Shuttle Rollout Forcing Function
James, George H., III; Carne, Thomas; Elliott, Kenny; Wilson, Bruce; [2005]; 11 pp.; In English; 23rd International Modal Analysis Conference and Expedition, 31 Jan. -3 Feb. 2005, Orlando, FL, USA; Copyright; Avail.: CASI: A03, Hardcopy
The Space Shuttle Vehicle is assembled in the Vertical Assembly Building (VAB) at Kennedy Space Flight Center in Florida. The Vehicle is stacked on a Mobile Launch Platform (MLP) that weighs eight million pounds. A Crawler Transporter (CT) then carries the MLP and the stacked vehicle (12 million pounds total weight) to the launch complex located 5 miles away. This operation is performed at 0.9 mph resulting in a 4.5-hour transport. A recent test was performed to monitor the dynamic environment that was produced during rollout. It was found that the rollout is a harmonic-rich dynamic environment that was previously not understood. This paper will describe work that has been performed to estimate the forcing function that is produced in the transportation process. The rollout analysis team has determined that there are two families of harmonics of the drive train, which excite the system as a function of CT speed. There are also excitation sources, which are random or narrow-band in frequency and are not a function of CT speed. This presentation will discuss the application of the Sum of Weighted Accelerations Technique (SWAT) to further refine this understanding by estimating the forces and moments at the center-of-mass. Author
Space Shuttles; Launching Pads; Launch Vehicle Configurations; Force Distribution
Source: NASA
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