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All sourcesNASA NTRSDOE OSTI
NASA NTRS1991

Mars sample return mission: Two alternate scenarios

Two scenarios for accomplishing a Mars Sample Return mission are presented. Mission A is a low-cost, low-mass scenario, while Mission B is a high-technology, high-science alternative. Mission A begins with the launch of one Titan 4 rocket with a Centaur G' upper stage. The Centaur performs the trans-Mars injection burn and is then released. The payload consists of two lander packages and the orbital transfer vehicle, which is responsible for supporting the landers during launch and interplanetary cruise. Near Mars, the landers separate - one bound for a polar site and the other for an equatorial site. After descending to the surface, the landers deploy small, local rovers to collect samples. The rovers return these samples to the landers for loading on the direct return rockets, which return the samples directly to the Earth's surface. Mission B starts with four Titan 4 launches, used to place the components of the planetary transfer vehicle (PTV) into orbit. The fourth launch payload is able to move to assemble the entire vehicle by simple docking routines. Once complete, the PTV begins a low-thrust trajectory out from low Earth orbit, through interplanetary space, and into low Mars orbit. It deploys a communications satellite into a one half sol orbit and then releases the lander pad cage at 500 km altitude. The lander package contains the lander, the Mars ascent vehicle (MAV), two lighter-than-air rovers (called Aereons), and one conventional land rover. The entire package is contained within a biconic aeroshell. After release from the PTV the lander package descends to the surface, where all three rovers are released to collect samples and map the terrain. The Aereons attempt to circumnavigate Mars and collect samples from a wide variety of sites, while the land rover examines a local area more thoroughly. The Aereons are equipped with small sample return rockets that can return their samples to the lander in the event that an Aereon is incapable of returning to the lander itself. Once all samples have been collected they are loaded onto the MAV and launched into orbit. The PTV then collects the samples and returns them to Earth orbit for recovery.

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NASA NTRSDate not supplied

NASA-ESA Mars Sample Return Program

NASA's Perseverance mission arrived at Jezero Crater on Mars in February 2021 and began scientific studies and acquisition of Martian samples for return to Earth by future missions, consistent with the recommendations of the U.S. science community in the previous Planetary Science Decadal Survey. NASA and ESA have established a joint Mars Sample Return (MSR) program to safely deliver these samples back to Earth, allowing researchers to use advanced scientific instrumentation that cannot be transported on robotic spacecraft and enable future studies of carefully curated samples using capabilities that have not yet been developed. The MSR architecture consists of two flight elements to follow Perseverance, the NASA-led Sample Retrieval Lander (SRL) and the ESA-led Earth Return Orbiter (ERO). The ERO is designed to orbit Mars and provide relay services for the SRL, including its ESA Sample Fetch Rover (SFR) and the NASA Mars Ascent Vehicle (MAV). The SRL deploys the SFR to retrieve Martian samples cached by the Perseverance rover and then returns the samples to the Orbiting Sample container (OS) on board the MAV using the ESA Sample Transfer Arm (STA). Independently, Perseverance could also deliver samples retained onboard to the OS. The MAV would launch and release the OS into low Mars orbit for rendezvous with the ERO. Upon successful capture of the OS in the ERO’s primary payload, the NASA Capture/Containment Return System (CCRS), the OS would be safely contained and loaded into the Earth Entry System (EES). The ERO will leave Mars orbit and release the EES on Earth approach on a ballistic reentry trajectory through the Earth's atmosphere for landing in the United States. Following return of the samples to Earth, the samples would be protected, preserved, assessed, curated, and made available to the international science community for scientific research and analysis. The NASA SRL and ESA ERO missions are expected to launch as early as 2026, with the return of Martian samples to Earth as early as 2031. MSR’s primary objective is the return of scientifically selected Mars samples for detailed investigation in terrestrial laboratories. The mission would also further inform the design of future human missions. The Mars Sample Return campaign is underway with the successful collection of several scientifically selected samples in Jezero Crater. The MSR Program is working towards a confirmation review in 2023 for the remaining flight elements.

Mars↗
NASA NTRS2001

STS-102 Countdown Status

George Diller, NASA Public Affairs, introduces Jeff Spaulding, NASA Test Director, Glenn Chin, Leonardo Payload Manager, and Ed Priselac, Shuttle Weather Officer, in this STS-102 prelaunch press conference. Mr. Spaulding gives an overview of the status of the Discovery Orbiter, including the prelaunch procedures (payload inspection and closure, avionics check, and the loading of the onboard cyrogenic tanks), the countdown and built in time holds, launch window, Discovery launch, and the landing. Mr. Chin discusses the payload status, including specifications on the Multi-Purpose Logistics Module, and the hardware contamination inspection. Mr. Priselac describes the weather forecast for the upcoming launch. The men then answer questions from the press.

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NASA NTRS2005

STS-114: Discovery L-2 Countdown Status Briefing

George Diller of NASA Public Affairs hosted this briefing. Pete Nickolenko, NASA Test Director; Scott Higgenbotham, STS-114 Payload-Mission Manager; Cathy Winters, Shuttle Weather Officer were present. Pete reports his team has completed the avionics system check ups, servicing of the cryogenic tanks will take about seven hours that day, and will perform engine system checks and pad close outs come evening. Pete also summarized other standard close out activities: check ups of the Orbiter and ground communications network, rotary service, structure retraction, and external tank load (ETL). Pete reported that the mission will be 12 days with two weather contingency days, and end of mission landing scheduled at Kennedy Space Center (KSC) at approximately 11:00 in the morning, Eastern time on July 25th. Scott briefly reported that all hardware is on board Discovery, closed out, and ready to fly. Cathy reported that hurricane Dennis moved to the North and looking forward to launch. She mentioned of a new hurricane looming and will be named Emily, spotted some crosswinds which will migrate to the west, there is 30% probability weather prohibiting launch. Cathy further gave current weather forecast supported with charts: the Launch Forecast, Tanking Forecast, SRB (Shuttle Solid Rocket Booster) Forecast, CONUS and TAL Launch Sites Forecast, and with 24 hours and 48 hours turn around plan. Launch constraints, weather, crosswinds, cloud cover, ground imagery system, launch countdown, launch crews, mission management simulations, launch team simulations were topics covered with the News Media.

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NASA NTRS1973

Graphite-epoxy door development for application to Space Shuttle.

The results of development of an advanced composite door landing-gear assembly are summarized. The development program consisted of design, analysis, fabrication, and structural testing of a door sector assembly designed to the operating loads expected for the Phase B Space Shuttle orbiter.

Kam, C. Y.↗
NASA NTRS1961

Estimates of Minimum Energy Requirements for Range-Controlled Return of a Nonlifting Satellite from a Circular Orbit

Existing expressions are used to obtain the minimum propellant fraction required for return from a circular orbit as a function of vacuum trajectory range. trajectory are matched to those of the atmospheric trajectory to obtain a complete return from orbit to earth. The results are restricted by the assumptions of (1) impulsive velocity change, (2) nearly circular transfer trajectory, ( 3) spherical earth, atmosphere, and gravitational field, (4) exponential atmospheric density variation with attitude, and (5) a nonrotating atmosphere. The solutions for the parameters of the vacuum Calculations are made t o determine the effects of longitudinal and lateral range on required propeUant fraction and reentry loading for a nonrotating earth and for several orbital altitudes. the single- and two-impulse method of return is made and the results indicate a "trade off" between propellant fraction required and landing- position accuracy. A comparison of An example of a return mission from a polar orbit is discussed where the initial deorbit point is the intersection of the North Pole horizon with the satellite orbit. Some effects of a rotating earth are also considered. It is found that, for each target-orbital-plane longitudinal difference, there exists a target latitude for which the required propellant fraction is a minimum.

Jackson, Charlie M., Jr.↗
NASA NTRS2007

Lunar Lander Structural Design Studies at NASA Langley

The National Aeronautics and Space Administration is currently developing mission architectures, vehicle concepts and flight hardware to support the planned human return to the Moon. During Phase II of the 2006 Lunar Lander Preparatory Study, a team from the Langley Research Center was tasked with developing and refining two proposed Lander concepts. The Descent-Assisted, Split Habitat Lander concept uses a disposable braking stage to perform the lunar orbit insertion maneuver and most of the descent from lunar orbit to the surface. The second concept, the Cargo Star Horizontal Lander, carries ascent loads along its longitudinal axis, and is then rotated in flight so that its main engines (mounted perpendicular to the vehicle longitudinal axis) are correctly oriented for lunar orbit insertion and a horizontal landing. Both Landers have separate crew transport volumes and habitats for surface operations, and allow placement of large cargo elements very close to the lunar surface. As part of this study, lightweight, efficient structural configurations for these spacecraft were proposed and evaluated. Vehicle structural configurations were first developed, and preliminary structural sizing was then performed using finite element-based methods. Results of selected structural design and trade studies performed during this activity are presented and discussed.

Wu, K. Chauncey↗
NASA NTRS1992

Design Feasibility Study of a Space Station Freedom Truss

Here, the focus is on the design and configuration feasibility of the short spacer for the Space Station Program in its launch configuration. The product of this study is being used by Rockwell International (Rocketdyne Division) as they continue their design concept of the current short spacer configuration. It is anticipated that the launch loads will dominate the on-orbit loads and dictate the design configuration of the short spacer. At the present time, the on-orbit loads have not been generated. The structural analysis discussed herein is based on the transient events derived from the Space Transportation System (STS) Interface Control Document (ICD). The transient loading events consist of liftoff loads, landing loads, and emergency landing loads. The quasi-static loading events have been neglected, since the magnitude of the acceleration factors are lower than the transient acceleration factors. The normal mode analyses presented herein are based on the most feasible configurations with acceptable stress ranges.

Armand, Sasan C.↗
NASA NTRS1982

The use of the Space Shuttle for land remote sensing

The use of the Space Shuttle for land remote sensing will grow significantly during the 1980's. The main use will be for general land cover and geological mapping purposes by worldwide users employing specialized sensors such as: high resolution film systems, synthetic aperture radars, and multispectral visible/IR electronic linear array scanners. Because these type sensors have low Space Shuttle load factors, the user's preference will be for shared flights. With this strong preference and given the present prognosis for Space Shuttle flight frequency as a function of orbit inclination, the strongest demand will be for 57 deg orbits. However, significant use will be made of lower inclination orbits. Compared with freeflying satellites, Space Shuttle mission investment requirements will be significantly lower. The use of the Space Shuttle for testing R and D land remote sensors will replace the free-flying satellites for most test programs.

Thome, P. G.↗
NASA NTRS2020

Challenges of Mars Sample Return Lander Entry, Descent, and Landing

The proposed Mars Sample Return (MSR) campaign would be perhaps the most ambitious robotic mission ever attempted in space exploration. The notional cam-paign consists of three Flagship-class missions operating in cooperation for over a decade in order to return samples of the Martian surface and atmosphere to Earth for analysis. The Mars 2020 rover, scheduled to launch in July 2020, will cache samples and place them on the surface for possible return. The second mission would be a Sample Return Lander (SRL) that consists of a small Sample Fetch Rover (SFR) to gather the samples, a Sample Transfer Arm (STA) to load the samples into a Mars Ascent Vehicle (MAV), and the MAV itself to launch the samples into orbit around Mars. The third mission would be an Earth Return Or-biter (ERO) designed to rendezvous and capture the Orbiting Sample (OS), return to Earth, and separate the Earth Entry Vehicle (EEV) for Entry, Descent, and Landing (EDL) at a location to be determined. This paper will focus on the SRL mission concept, specifically the EDL phase. Given the ambitious SRL sample re-trieval baseline surface mission, including a rocket launch of the samples into Mars orbit, it is estimated that the EDL system may be required to deliver as much as 2100 kg of dry mass to the surface. This represents an approximate 20-25% in-crease in mass capability over previous landed Mars missions. Additionally, there is a high probability that SRL would have to land very close to the samples on the surface to expedite retrieval operations; therefore, Pin Point Landing (PPL) accu-racy may be required. To address these challenges, promising EDL configuration augmentations were studied to include larger forebody/higher drag entry capsules, hypersonic/supersonic inflatable/non-inflatable aerodynamic decelerators, hypersonic trim tabs, ballute drag devices, larger parachutes, higher Mach and higher dynamic pressure parachute deployments, lower parachute deployment altitudes having shorter chute timelines necessitating more efficient terrain sensor strategies, and ad-ditional fuel for longer powered descent diverts to the target landing site. Over-arching the entire trade study was an attempt to stay as close to the experience base of past successful missions as possible to reduce implementation cost and risk. This paper will discuss the entire SRL EDL trade study in detail. The information presented about the potential MSR campaign is pre-decisional and is provided for planning and discussion purposes only.

Ivanov, Mark C.↗
NASA NTRS1981

The long duration exposure facility structural interface

The development of an efficient structural interface between the Long Duration Exposure Facility (LDEF) and the space shuttle orbiter is discussed. Thermal gradients and manufacturing tolerances restrict the number of retention points in the shuttle to four with no more than three in one plane. Loads during the powered exit, high drag reentry, and landing are also critical factors in the location of retention points. A LDEF design is described that meets all requirements while avoiding the structural weight penalties associated with an asymmetrical arrangement. The ground handling and assembly problems are resolved in an economical and effective manner.

Long, M. J.↗
NASA NTRS2004

Space Shuttle Crawler Transporter Vibration Analysis in Support of Rollout Fatigue Load Spectra Verification Program

The Crawler Transporter (CT), designed and built for the Apollo Program in the 1960's and surpassing its initial operational life, has become an integral part of the Space Shuttle Program (SSP). The CT transports the Space Shuttle Vehicle (SSV) stack, atop the Mobile Launch Platform (MLP), from the Vehicle Assembly Building (VAB) to the launch pad. This support structure provides hydraulic jacking, leveling and load equalization for the 12 million pound stack on its 3.5-5.0 mile rollout to the launch pad. Major elements of the SSV, consisting of the orbiter, solid rocket boosters (SRB) and external tank (ET) have required fatigue analyses as part of the mission life certification. Compared to rollout vibration, the SSV sees relatively high vibration loads during launch, ascent, descent and landing phases of the mission. Although preliminary measured SRB vibration levels during rollout were of low amplitude and frequency, the duration of the rollout phase is typically high, from 5-6 hours. As part of an expanded mission life assessment, additional certification effort was initiated to define fatigue load spectra for rollout. This study addresses the CT vibration analyses in support of the rollout fatigue study. Structural models developed for modal and vibration analyses were used to identify unique CT, CT/MLP and CT/MLP/SRB vibration characteristics for comparison to instrumented rollout tests. Whereas the main structural and vibration characteristics of the SSV are well defined, minimum analytical and vibration test data on the Crawler Transporter were available. Unique vibration characteristics of the CT are attributable to the drive mechanism, hydraulic jacking system, structural framing and the CT-to-MLP support pad restraints. Initial tests performed on the CT/MLP/SRB configuration showed reasonable correlation with predicted mode shapes and frequencies.

Margasahayam, Ravi N.↗
NASA NTRS2006

A Low Cost Spacecraft Architecture for Robotic Lunar Exploration Projects

A program of frequent, capable, but affordable lunar robotic missions prior to return of humans to the moon can contribute to the Vision for Space Exploration (VSE) NASA is tasked to execute. The Lunar Reconnaissance Orbiter (LRO) and its secondary payload are scheduled to orbit the moon, and impact it, respectively, in 2008. It is expected that the sequence of missions occurring for approximately the decade after 2008 will place an increasing emphasis on soft landed payloads. These missions are requited to explore intrinsic characteristics of the moon, such as hydrogen distribution in the regolith, and levitated dust, to demonstrate the ability to access and process in-situ resources, and to demonstrate functions critical to supporting human presence, such as automated precision navigation and landing. Additional factors governing the design of spacecraft to accomplish this diverse set of objectives are: operating within a relatively modest funding profile, the need tb visit multiple sites (both polar and equatorial) repeatedly, and to use the current generation of launch vehicles. In the US, this implies use of the Evolved Expendable Launch Vehicles, or EELVs, although this design philosophy may be extended to launch vehicles of other nations, as well. Many of these factors are seemingly inconsistent with each other. For example, the cost of a spacecraft usually increases with mass; therefore the desire to fly frequent, modestly priced spacecraft seems to imply small spacecraft (< 1 Mt, injected mass). On the other hand, the smallest of the EELVs will inject approx. 3 Mt. on a Trans Lunar Injection (TLI) trajectory md would therefore be wasteful or launching a single, small spacecraft. Increasing the technical capability of a spacecraft (such as autonomous navigation and soft landing) also usually increases cost. A strategy for spacecraft design that meets these conflicting requirements is presented. Taken together, spacecraft structure and propulsion subsystems constitute the majority of spacecraft mass; saving development and integration cost on these elements is critical to controlling cost. Therefore, a low cost, modular design for spacecraft structure and propulsion subsystems is presented which may be easily scaled up or down for either insertion into lunar orbit or braking for landing on the lunar surface. In order to effectively use the approx.3 Mt mass-to-TLI of the EELV, two low cost spacecraft will be manifested on the same launch. One spacecraft will be located on top of the other for launch and the two will have to be released in sequence in order to achieve all mission objectives. The two spacecraft could both be landers, both orbiters, or one lander and one orbiter. In order to achieve mass efficiency, the body of the spacecraft will serve the dual purposes of carrying launch loads and providing attachment points for all the spacecraft subsystems. In order to avoid unaffordable technology development costs, small liquid propulsion components and autonomous, scene-matching navigation cameras may be adapted from military missile programs in order to execute precision soft landings.

Lemke, Lawrence G.↗
NASA NTRS2012

Plasticity of Neurovestibular Systems Following Micro- and Hyper-Gravity Exposure and Readaptation to Earth's 1G

The gravity-sensing organs sense the sum of inertial force due to head translation and head orientation relative to gravity. Normally gravity is constant, and yet the neural sensors show remarkable plasticity. When the force of gravity changes, such as in spaceflight or during centrifugation, the neurovestibular system responds by regulating its neural output, and this response is similar for the vertebrate utricular nerve afferents and for the statocyst hair cell in invertebrates. First, we examine the response of utricular afferents in toadfish following exposure to G on two orbital missions (STS-90 and 95). Within the first day after landing, magnitude of neural response to an applied acceleration was significantly elevated, and re-adaptation back to control values occurred within approximately 30 hours. Time course of return to normal approximately parallels the decrease in vestibular disorientation in astronauts following return. Next, we use well-controlled hyper-G experiments in the vertebrate model to address: If G leads to adaptation and subsequent re-adaptation neural processes, does the transfer from 1G to hyper-G impart the opposite effects and do the effects accompanying transfer from the hyper-G back to the 1G conditions resemble as an analog the transfer from 1G to the microG Results show a biphasic pattern in reaction to 3G exposures: an initial sensitivity up-regulation (3- and 4-day) followed by a significant decrease after longer exposure. Return to control values is on the order of 4-8 days. Utricular sensitivity is strongly regulated up or down by gravity load and the duration of exposure. Interestingly, we found no correlation of response and hair cell synaptic body counts despite the large gain difference between 4- and 16-Day subjects. Lastly, we examine responses of statocyst receptors in land snail following exposure to G on two unmanned Russian Orbital missions (Foton M-2 and -3). Here, we have the ability to measure the output directly from the hair cells. Similar to afferents in vertebrates the hair cells increased their response sensitivity to vestibular stimulation. Two major pieces of information are needed: the precise vertebrate hair cell response to altered gravity and the impact of longer duration exposures on sensory plasticity.

Boyle, Richard D.↗
NASA NTRS2013

NASA Tech Briefs, September 2013

Topics include: ISS Ammonia Leak Detection Through X-Ray Fluorescence; A System for Measuring the Sway of the Vehicle Assembly Building; Fast, High-Precision Readout Circuit for Detector Arrays; Victim Simulator for Victim Detection Radar; Hydrometeor Size Distribution Measurements by Imaging the Attenuation of a Laser Spot; Quasi-Linear Circuit; High-Speed, High-Resolution Time-to-Digital Conversion; Li-Ion Battery and Supercapacitor Hybrid Design for Long Extravehicular Activities; Ultrasonic Low-Friction Containment Plate for Thermal and Ultrasonic Stir Weld Processes; High-Powered, Ultrasonically Assisted Thermal Stir Welding; Next-Generation MKIII Lightweight HUT/Hatch Assembly; Centrifugal Sieve for Gravity-Level-Independent Size; Segregation of Granular Materials; Ion Exchange Technology Development in Support of the Urine Processor Assembly; Nickel-Graphite Composite Compliant Interface and/or Hot Shoe Material; UltraSail CubeSat Solar Sail Flight Experiment; Mechanism for Deploying a Long, Thin-Film Antenna From a Rover; Counterflow Regolith Heat Exchanger; Acquisition and Retaining Granular Samples via a Rotating Coring Bit; Very-Low-Cost, Rugged Vacuum System; Medicine Delivery Device With Integrated Sterilization and Detection; FRET-Aptamer Assays for Bone Marker Assessment, C-Telopeptide, Creatinine, and Vitamin D; Multimode Directional Coupler for Utilization of Harmonic Frequencies from TWTAs; Dual-Polarization, Multi-Frequency Antenna Array for use with Hurricane Imaging Radiometer; Complementary Barrier Infrared Detector (CBIRD) Contact Methods; Autonomous Control of Space Nuclear Reactors; High-Power, High-Speed Electro-Optic Pockels Cell Modulator; Covariance Analysis Tool (G-CAT) for Computing Ascent, Descent, and Landing Errors; Enigma Version 12; Micrometeoroid and Orbital Debris (MMOD) Shield Ballistic Limit Analysis Program; Spitzer Telemetry Processing System; Planetary Protection Bioburden Analysis Program; Wing Leading Edge RCC Rapid Response Damage Prediction Tool (IMPACT2); ISSM: Ice Sheet System Model; Automated Loads Analysis System (ATLAS); Integrated Main Propulsion System Performance Reconstruction Process/Models. Phoenix Telemetry Processor; Contact Graph Routing Enhancements Developed in ION for DTN; GFEChutes Lo-Fi; Advanced Strategic and Tactical Relay Request Management for the Mars Relay Operations Service; Software for Generating Troposphere Corrections for InSAR Using GPS and Weather Model Data; Ionospheric Specifications for SAR Interferometry (ISSI); Implementation of a Wavefront-Sensing Algorithm; Sally Ride EarthKAM - Automated Image Geo-Referencing Using Google Earth Web Plug-In; Trade Space Specification Tool (TSST) for Rapid Mission Architecture (Version 1.2); Acoustic Emission Analysis Applet (AEAA) Software; Memory-Efficient Onboard Rock Segmentation; Advanced Multimission Operations System (ATMO); Robot Sequencing and Visualization Program (RSVP); Automating Hyperspectral Data for Rapid Response in Volcanic Emergencies; Raster-Based Approach to Solar Pressure Modeling; Space Images for NASA JPL Android Version; Kinect Engineering with Learning (KEWL); Spacecraft 3D Augmented Reality Mobile App; MPST Software: grl_pef_check; Real-Time Multimission Event Notification System for Mars Relay; SIM_EXPLORE: Software for Directed Exploration of Complex Systems; Mobile Timekeeping Application Built on Reverse-Engineered JPL Infrastructure; Advanced Query and Data Mining Capabilities for MaROS; Jettison Engineering Trajectory Tool; MPST Software: grl_suppdoc; PredGuid+A: Orion Entry Guidance Modified for Aerocapture; Planning Coverage Campaigns for Mission Design and Analysis: CLASP for DESDynl; and Space Place Prime.

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NASA NTRS1981

Cornering characteristics of the nose-gear tire of the space shuttle orbiter

An experimental investigation was conducted to evaluate cornering characteristics of the 32 x 8.8 nose gear tire of the space shuttle orbiter. Data were obtained on a dry concrete runway at nominal ground speeds ranging from 50 to 100 knots and over a range of tire vertical loads and yaw angles which span the expected envelope of loads and yaw angles to be encountered during space shuttle landing operations. The cornering characteristics investigated included side and drag forces and friction coefficients, aligning and overturning torques, friction force moment arm, and the lateral center of pressure shift. Results of this investigation indicate that the cornering characteristics of the space shuttle nose gear tire are insensitive to variations in ground speed over the range tested. The effects on cornering characteristics of variations in the tire vertical load and yaw angle are as expected. Trends observed are consistent with trends observed during previous cornering tests involving other tire sizes.

Vogler, W. A.↗
NASA NTRS1988

Analysis of the Space Shuttle Orbiter skin panels under simulated hydrodynamic loads

The Space Shuttle orbiter skin panels were analyzed under pressure loads simulating hydrodynamic loads to determine their capability to sustain a potential ditching and to determine pressures that typically would produce failures. Two Dynamic Crash Analysis of Structures (DYCAST) finite element models were used. One model was used to represent skin panels (bays) in the center body, while a second model was used to analyze a fuselage bay in the wing region of the orbiter. From an assessment of the DYCAST nonlinear computer results, it is concluded that the probability is extremely high that most, if not all, of the lower skin panels would rupture under ditching conditions. Extremely high pressure loads which are produced under hydrodynamic planning conditions far exceed the very low predicted failure pressures for the skin panels. Consequently, a ditching of the orbiter is not considered to have a high probability of success and should not be considered a means of emergency landing unless no other option exists.

Carden, Huey D.↗
NASA NTRS1993

Orbiter lessons learned: A guide to future vehicle development

Topics addressed are: (1) wind persistence loads methodology; (2) emphasize supportability in design of reusable vehicles; (3) design for robustness; (4) improved aerodynamic environment prediction methods for complex vehicles; (5) automated integration of aerothermal, manufacturing, and structures analysis; (6) continued electronic documentation of structural design and analysis; and (7) landing gear rollout load simulations.

Greenberg, Harry Stan↗