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NASA NTRS2006

Planetary Dust: Cross-Functional Considerations

Apollo astronauts learned first hand how problems with dust impact lunar surface missions. After three days, lunar dust contaminating on EVA suit bearings led to such great difficulty in movement that another EVA would not have been possible. Dust clinging to EVA suits was transported into the Lunar Module. During the return trip to Earth, when microgravity was reestablished, the dust became airborne and floated through the cabin. Crews inhaled the dust and it irritated their eyes. Some mechanical systems aboard the spacecraft were damaged due to dust contamination. Study results obtained by Robotic Martian missions indicate that Martian surface soil is oxidative and reactive. Exposures to the reactive Martian dust will pose an even greater concern to the crew health and the integrity of the mechanical systems. As NASA embarks on planetary surface missions to support its Exploration Vision, the effects of these extraterrestrial dusts must be well understood and systems must be designed to operate reliably and protect the crew in the dusty environments of the Moon and Mars. The AIM Dust Assessment Team was tasked to identify systems that will be affected by the respective dust, how they will be affected, associated risks of dust exposure, requirements that will need to be developed, identified knowledge gaps, and recommended scientific measurements to obtain information needed to develop requirements, and design and manufacture the surface systems that will support crew habitation in the lunar and Martian outposts.

Wagner, Sandra↗

NASA NTRS1973

Apollo LM guidance computer software for the final lunar descent.

In all manned lunar landings to date, the lunar module Commander has taken partial manual control of the spacecraft during the final stage of the descent, below roughly 500 ft altitude. This report describes programs developed at the Charles Stark Draper Laboratory, MIT, for use in the LM's guidance computer during the final descent. At this time computational demands on the on-board computer are at a maximum, and particularly close interaction with the crew is necessary. The emphasis is on the design of the computer software rather than on justification of the particular guidance algorithms employed. After the computer and the mission have been introduced, the current configuration of the final landing programs and an advanced version developed experimentally by the author are described.

Eyles, D.↗

NASA NTRS2004

An Assessment of Dust Effects on Planetary Surface Systems to Support Exploration Requirements

Apollo astronauts learned first hand how problems with dust impact lunar surface missions. After three days, lunar dust contamination on EVA suit bearings led to such great difficulty in movement that another EVA would not have been possible. Dust clinging to EVA suits was transported into the Lunar Module. During the return trip to Earth, when micro gravity was reestablished, the dust became airborne and floated through the cabin. Crews inhaled the dust and it irritated their eyes. Some mechanical systems aboard the spacecraft were damaged due to dust contamination. Study results obtained by Robotic Martian missions indicate that Martian surface soil is oxidative and reactive. Exposures to the reactive Martian dust will pose an even greater concern to the crew health and the integrity of the mechanical systems. As NASA embarks on planetary surface missions to support its Exploration Vision, the effects of these extraterrestrial dusts must be well understood and systems must be designed to operate reliably and protect the crew in the dusty environments of the Moon and Mars. The AIM Dust Assessment Team was tasked to identify systems that will be affected by the respective dust, how they will be affected, associated risks of dust exposure, requirements that will need to be developed, identified knowledge gaps, and recommended scientific measurements to obtain information needed to develop requirements, and design and manufacture the surface systems that will support crew habitation in the lunar and Martian outposts.

Wagner, Sandy↗

NASA NTRS2009

Ares V: Game Changer for National Security Launch

NASA is designing the Ares V cargo launch vehicle to vastly expand exploration of the Moon begun in the Apollo program and enable the exploration of Mars and beyond. As the largest launcher in history, Ares V also represents a national asset offering unprecedented opportunities for new science, national security, and commercial missions of unmatched size and scope. The Ares V is the heavy-lift component of NASA's dual-launch architecture that will replace the current space shuttle fleet, complete the International Space Station, and establish a permanent human presence on the Moon as a stepping-stone to destinations beyond. During extensive independent and internal architecture and vehicle trade studies as part of the Exploration Systems Architecture Study (ESAS), NASA selected the Ares I crew launch vehicle and the Ares V to support future exploration. The smaller Ares I will launch the Orion crew exploration vehicle with four to six astronauts into orbit. The Ares V is designed to carry the Altair lunar lander into orbit, rendezvous with Orion, and send the mated spacecraft toward lunar orbit. The Ares V will be the largest and most powerful launch vehicle in history, providing unprecedented payload mass and volume to establish a permanent lunar outpost and explore significantly more of the lunar surface than was done during the Apollo missions. The Ares V consists of a Core Stage, two Reusable Solid Rocket Boosters (RSRBs), Earth Departure Stage (EDS), and a payload shroud. For lunar missions, the shroud would cover the Lunar Surface Access Module (LSAM). The Ares V Core Stage is 33 feet in diameter and 212 feet in length, making it the largest rocket stage ever built. It is the same diameter as the Saturn V first stage, the S-IC. However, its length is about the same as the combined length of the Saturn V first and second stages. The Core Stage uses a cluster of five Pratt & Whitney Rocketdyne RS-68B rocket engines, each supplying about 700,000 pounds of thrust. Its propellants are liquid hydrogen and liquid oxygen. The two solid rocket boosters provide about 3.5 million pounds of thrust at liftoff. These 5.5-segment boosters are derived from the 4-segment boosters now used on the Space Shuttle, and are similar to those used in the Ares I first stage. The EDS is powered by one J-2X engine. The J-2X, which has roughly 294,000 pounds of thrust, also powers the Ares I Upper Stage. It is derived from the J-2 that powered the Saturn V second and third stages. The EDS performs two functions. Its initial suborbital burns will place the lunar lander into a stable Earth orbit. After the Orion crew vehicle, launched separately on an Ares I, docks with the lander/EDS stack, EDS will ignite a second time to put the combined 65-metric ton vehicle into a lunar transfer orbit. When it stands on the launch pad at Kennedy Space Center late in the next decade, the Ares V stack will be approximately 381 feet tall and have a gross liftoff mass of 8.1 million pounds. The current point-of-departure design exceeds Saturn V s mass capability by approximately 40 percent. Using the current payload shroud design, Ares V can carry 315,000 pounds to 29-degree low Earth orbit (LEO) or 77,000 pounds to a geosynchronous orbit. Another unique aspect of the Ares V is the 33-foot-diameter payload shroud, which encloses approximately 30,400 cubic feet of usable volume. A larger hypothetical shroud for encapsulating larger payloads has been studied. While Ares V makes possible larger payload masses and volumes, it may alternately make possible more cost-effective mission design if the relevant payload communities are willing to consider an alternative to the existing approach that has driven them to employ complexity to solve current launch vehicle mass and volume constraints. By using Ares V s mass and volume capabilities as margin, payload designers stand to reduce development risk and cost. Significant progress has been made on the Ares V to support a plaed fiscal 2011 authority-to-proceed (ATP) milestone. The Ares V team is actively reaching out to external organizations during this early concept phase to ensure that the Ares V vehicle can be leveraged for national security, science, and commercial development needs. This presentation will discuss Ares V vehicle configuration, the path to the current concept, accomplishments to date, and potential payload utilization opportunities.

Sumrall, Phil↗

NASA NTRS2009

A Review of the MLAS Parachute Systems

The NASA Engineering and Safety Center (NESC) is developing the Max Launch Abort System (MLAS) as a risk-mitigation design should problems arise with the baseline Orion spacecraft launch abort design. The Max in MLAS is dedicated to Max Faget, the renowned NASA spacecraft designer. The MLAS flight test vehicle consists of boost skirt, coast skirt and the MLAS fairing which houses a full scale boilerplate Orion Crew Module (CM). The objective of the flight test is to prove that the CM can be released from the MLAS fairing during pad abort conditions without detrimental recontact between the CM and fairing, achieving performance similar to the Orion launch abort system. The boost and coast skirts provide the necessary thrust and stability to achieve the flight test conditions and are released prior to the test -- much like the Little Joe booster was used in the Apollo Launch Escape System tests. To achieve the test objective, two parachutes are deployed from the fairing to reorient the CM/fairing to a heatshield first orientation. The parachutes then provide the force necessary to reduce the total angle of attack and body angular rates required for safe release of the CM from the fairing. A secondary test objective after CM release from the fairing is to investigate the removal of the CM forward bay cover (FBC) with CM drogue parachutes for the purpose of attempting to synchronously deploying a set of CM main parachutes. Although multiple parachute deployments are used in the MLAS flight test vehicle to complete its objective, there are only two parachute types employed in the flight test. Five of the nine parachutes used for MLAS are 27.6 ft D(sub 0) ribbon parachutes, and the remaining four are standard G-12 cargo parachutes. This paper presents an overview of the 27.6 ft D(sub 0) ribbon parachute system employed on the MLAS flight test vehicle for coast skirt separation, fairing reorientation, and as drogue parachutes for the CM after separation from the fairing. Discussion will include: the process used to select this design, previously proven as a spin/stall recovery parachute; descriptions of all components of the parachute system; the minor modifications necessary to adapt the parachute to the MLAS program; the techniques used to analyze the parachute for the multiple roles it performs; a discussion of the rigging techniques used to interface the parachute system to the vehicle; and a brief description of how the evolution of the program affected parachute usage and analysis. An overview of the Objective system, rationale for the MLAS approach and the future of the program will also be presented. We hope to have flight test results to report at the time of the Conference Presentation.

Taylor, Anthony P.↗

NASA NTRS1994

Apollo 11 Recovery

The Apollo 11 spacecraft is lifted from the ocean after its water landing (not shown) onto the deck of the USS Hornet. The crewmembers, Commander Neil A. Armstrong, Lunar Module Pilot Edwin E. Aldrin, Jr., and Command Module Pilot Michael Collins, are seen going from the spacecraft to the quarantine chambers. They talk to family members over the phone after their arrival at the Ellington Air Force Base. Footage shows the celebration of Armstrong's birthday and the release of the crew from quarantine.

Source record↗

NASA NTRS1992

Geology of the Apollo 17 site

The Apollo 17 landing site was unique in several aspects: (1) it was the only site that was not selected from telescopic-based geologic interpretation--interest in the site was generated by the visual observations of Al Worden, Apollo 15 Command Module pilot, who interpreted dark-haloed craters as possible cinder cones; (2) instead of 20-m-resolution photographs, as was the norm for all earlier missions, this site had Apollo 15 panoramic camera photography coverage that had 2-m resolution; and (3) it had a geologist-astronaut aboard who was intimately involved in all stages of planning and mission operation, and was also instrumental in the design of a long-handled sample bag holder that eliminated the need for crew to dismount before collecting a sample, which then permitted sampling between major stations. Details of site geology, sample description, and geologic synthesis of the site as viewed from studies through 1976 are summarized.

Muehlberger, W. R.↗

NASA NTRS1995

Smoke detection in low-G fires

Fires in spacecraft are considered a credible risk. To respond to this risk, NASA flew fire detectors on Skylab and the Space Shuttle (STS) and included them in the design for International Space Station Alpha (ISSA). In previous missions (Mercury, Gemini and Apollo), the crew quarters were so cramped that it was not considered credible that the astronauts could fail to observe a fire. The Skylab nodule included approximately 20 UV fire detectors. The space shuttle has 9 ionization detectors in the mid deck and flight deck and Spacelab has six additional ionization detectors. The planned detectors for ISSA are laser-diode, forward-scattering, smoke or particulate detectors. Current plans for the ISSA call for two detectors in the open area of the module and detectors in racks that have both cooling air flow and electrical power. Due to the complete absence of data concerning the nature of particulate and radiant emission from low-g fires, all three of these detector systems were designed based upon 1-g test data. As planned mission durations and complexity increase and the volume of spacecraft increases, the need for and importance of effective, crew independent, fire detection grows significantly. This requires more knowledge concerning low-gravity fires and how they might be detected. To date, no combustion-generated particulate samples have been collected for well-developed microgravity flames. All of the extant data come from drop tower tests and therefore only correspond to the early stages of a fire. The fuel sources were restricted to laminar gas-jet diffusion flames and rapidly overheated wire insulation. These gas-jet drop tower tests indicate, through thermophoretic sampling, that soot primaries and aggregates (groups of primary particles) in micro-g may be significantly larger than those in normal-g (ng). This raises new scientific questions about soot processes as well as practical issues for particulate detection/alarm threshold levels used in on-orbit smoke detectors. Furthermore, it is widely speculated but unverified that the aggregates will grow to very large scales in a microgravity fire of longer duration than available on the ground. Preliminary tests in the 2.2 second drop tower suggest that particulate generated by overheated wire insulation will also be larger in microgravity than in normal gravity. TEM grids downstream of the fire region in the WIF experiment as well as visual observation of long string-like aggregates, further confirm this suggestion. The combined impact of these limited results and theoretical predictions is that direct knowledge of low-g combustion particulate as opposed to extrapolation from 1-g data is needed for a more confident design of smoke detectors for spacecraft.

Urban, David L.↗

NASA NTRS1994

Comment on 'Guidance for an Aerocapture Maneuver

It is stated that the aerodynamic forces on the vehicle being aerocaptured are controlled by "altering the angle of attack" and thereby controlling the lift coefficient. Furthermore, the resulting variation of drag coefficient with angle of attack was ignored. The purpose of this Comment is to point out that an aerodynamic control method that is much more effective than the pitch modulation has been studied and utilized during entries for many years. During aerocapture, it is desirable to have a large range of lift coefficients available, while keeping the vehicle's ballistic coefficients constant. This is accomplished by modulating the vehicle's bank angle, i.e., by rolling the vehicle about its velocity vector. By this method, the angle of attack can be held constant (at the trim angle, if desired), and the C(sub D) and the ballistic coefficient remain constant. Furthermore, the vertical component of the normal force vector (essentially the lift) can be varied over its entire range, from maximum positive to maximum negative values. Reaction controls, rather than aerodynamic ones, are usually utilized to change the bank angle of the vehicle, thus requiring the use of fuel. However, the fuel expenditure that is required to change the bank angle is far less than the amount that would have to be used to continuously hold the vehicle at pitch angles that differ significantly from its trim angle of attack. Also, it has been shown that bank angle modulation to vary the lift can enlarge the entry corridor by increasing the entry angle for the undershoot boundary, where both the heating rate and deceleration reach a maximum. Finally, the crew's deceleration tolerance can be increased somewhat when the bank angle is varied, as opposed to the pitch angle. For bank modulation, the deceleration force vector can be kept at a constant angle with respect to the occupants whose tolerance to g loads is highest when the force is applied in a direction normal to the upper torso. The advantages of bank angle variation to modulate the lift vector were recognized long ago, and this method of control was used successfully on the Apollo command module during lunar return' and, more recently, for the Space Shuttle Orbiter.

Tauber, Michael E.↗

NASA NTRS2005

Analysis of Heating Rates on the Conical Surface of Apollo Command Module Flying AS-202 Flight

The aerodynamic and aerothermal heating at the leeward surfaces of the Apollo capsule flying high angle of attack were found difficult to simulate using conventional tools. Due to the large subsonic region in the proximity of the shoulder of the base shield, correlation-based tools were found to be inadequate. CFD tools are too time consuming for conceptual design purposes, and cannot account for the transient effects of material response such as wall temperature and blowing. An accurate and timely simulation is essential to effectively size the thermal protection system (TPS), to enhance its performance, and to ensure the safety of the crew. Northrop Grumman Corporation applied a modified version of the MASCC/ATAC program, an inviscid flowfield code with boundary layer solver to simulate this scenario. The MASCC/ATAC code is believed to be the only non-CFD code that can rigorously perform the simulations on the windward surface. The predictions on the windward side of the conical surface were found to be in good agreement with flight data over a wide range of environments. The results are presented in the paper.

Wong, Jim L.↗

NASA NTRS2010

Orion Pad Abort 1 GN and C Design and Development

The first flight test of the Orion Abort Flight Test project is scheduled to launch in Spring 2010. This flight test is known as Pad Abort 1 (PA-1) and it is intended to accomplish a series of flight test objectives, including demonstrating the capability of the Launch Abort System (LAS) to propel the Crew Module (CM) to a safe distance from a launch vehicle during a pad abort. The PA-1 Flight Test Article (FTA) is actively controlled by a guidance, navigation, and control (GN&C) system for much of its flight. The purpose of this paper is to describe the design, development, and analysis of the PA-1 GN&C system. A description of the technical solutions that were developed to meet the challenge of satisfying many competing requirements is presented. A historical perspective of how the Orion LAV compares to the Apollo Launch Escape Vehicle (LEV) design will also be included.

Medina, Edgar A.↗

NASA NTRS2009

Apollo Spacecraft and Saturn V Launch Vehicle Pyrotechnics/Explosive Devices

The Apollo Mission employs more than 210 pyrotechnic devices per mission.These devices are either automatic of commanded from the Apollo spacecraft systems. All devices require high reliability and safety and most are classified as either crew safety critical or mission critical. Pyrotechnic devices have a wide variety of applications including: launch escape tower separation, separation rocket ignition, parachute deployment and release and electrical circuit opening and closing. This viewgraph presentation identifies critical performance, design requirements and safety measures used to ensure quality, reliability and performance of Apollo pyrotechnic/explosive devices. The major components and functions of a typical Apollo pyrotechnic/explosive device are listed and described (initiators, cartridge assemblies, detonators, core charges). The presentation also identifies the major locations and uses for the devices on: the Command and Service Module, Lunar Module and all stages of the launch vehicle.

Interbartolo, Michael↗

NASA NTRS1972

Preliminary geologic investigation of the Apollo 15 landing site

The Apollo 15 lunar module (LM) landed on the mare surface of Palus Putredinis on the eastern edge of the Imbrium Basin. The site is between the Apennine Mountain front and Hadley Rille. The objectives of the mission, in order of decreasing priority, were description and sampling of three major geologic features-the Apennine Front, Hadley Rille, and the mare. The greater number of periods of extravehicular activity (EVA) and the mobility provided by the lunar roving vehicle (ROVER) allowed much more geologic information to be obtained from a much larger area than those explored by previous Apollo crews. A total of 5 hours was spent at traverse station stops, and the astronauts transmitted excellent descriptions of the lunar surface while in transit between stations.

Swann, G. A.↗

NASA NTRS1973

Mission requirements: Skylab rescue mission SL-R

The Skylab Program includes three low earth orbit missions. These missions are designated SL-1/SL-2,SL-3 and SL-4. In addition to the three nominal Skylab missions, the program includes the Skylab Rescue Mission (SL-R). The SL-R mission is designed to provide a safe return of the Skylab crew in the event the Command Service Module (CSM) becomes disabled while docked to the Saturn Workshop (SWS). Mission requirements for the SL-R mission only are presented. SL-R mission configuration will be a CSM (modified with a field installed kit) manned by two crewmen launched on a Saturn IB Launch Vechicle. The SL-R CSM will rendezvous and dock with the SWS (or Orbital Assembly (OA), consisting of the SWS and disabled CSM, if the disabled CSM has not previously been jettisoned). The SWS configuration includes a Multiple Docking Adapter (MDA), Apollo Telescope Mount (ATM), Airlock Module (AM), and an S-IVB stage (modified as an Orbital Workshop (OWS), previously launched and inserted into orbit on a two-stage Saturn V Launch Vehicle for the SL-1/SL-2 mission.

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NASA NTRS2006

Airborne Dust in Space Vehicles and Habitats

Airborne dust, suspended inside a space vehicle or in future celestial habitats, can present a serious threat to crew health if it is not controlled. During the Apollo missions to the moon, lunar dust brought inside the capsule caused eye irritation and breathing difficulty to the crew when they launched from the moon and re-acquired "microgravity." During Shuttle flights reactive and toxic dusts such as lithium hydroxide have created a risk to crew health, and fine particles from combustion events can be especially worrisome. Under nominal spaceflight conditions, airborne dusts and particles tend to be larger than on earth because of the absence of gravity settling. Aboard the ISS, dusts are effectively managed by HEPA filters, although floating dust in newly-arrived modules can be a nuisance. Future missions to the moon and to Mars will present additional challenges because of the possibility that external dust will enter the breathing atmosphere of the habitat and reach the crew's respiratory system. Testing with simulated lunar and Martian dust has shown that these materials are toxic when placed into the lungs of test animals. Defining and evaluating the physical and chemical properties of Martian dusts through robotic missions will challenge our ability to prepare better dust simulants and to determine the risk to crew health from exposure to such dusts.

James, John↗

NASA NTRS1974

Apollo lunar descent guidance

Apollo lunar-descent guidance transfers the Lunar Module from a near-circular orbit to touchdown, traversing a 17 deg central angle and a 15 km altitude in 11 min. A group of interactive programs in an onboard computer guide the descent, controlling altitude and the descent propulsion system throttle. A ground-based program pre-computes guidance targets. The concepts involved in this guidance are described. Explicit and implicit guidance are discussed, guidance equations are derived, and the earlier Apollo explicit equation is shown to be an inferior special case of the later implicit equation. Interactive guidance, by which the two-man crew selects a landing site in favorable terrain and directs the trajectory there, is discussed. Interactive terminal-descent guidance enables the crew to control the essentially vertical descent rate in order to land in minimum time with safe contact speed. The altitude maneuver routine uses concepts that make gimbal lock inherently impossible.

Klumpp, A. R.↗

NASA NTRSDate not supplied

Trends in Human Spaceflight: Analysis of Observed Propulsion Failure Modes Following NASA’s Artemis I Mission

In 2022, NASA’s Artemis program returned the agency to lunar space, propelled by the Orion European Service Module (ESM). Through rigorous ground test campaigns for Artemis I and II and the collection and transmission of operational data during the Artemis I flight, the Safety and Mission Assurance (S&MA) team created and maintained a robust database of component nonconformances. This study analyzes in-flight and ground processing findings on the Artemis ESM propulsion subsystem to track reliability and safety metrics, as well as comparing the data to that of previous human spaceflight programs, Apollo and Space Shuttle. In drawing a comparison to the Artemis I flight, recommendations can be made for the Artemis II crewed mission to ensure best practices for flight safety, including redundancy and failure tolerance in the system. The results reinforce a need for robust safety standards and caution against the complacency of a single successful test flight.

human spaceflight safety↗

NASA NTRSDate not supplied

Trends in Human Spaceflight: Analysis of Observed Propulsion Failure Modes Following NASA’s Artemis I Mission

human spaceflight safety↗