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NASA NTRSDate not supplied

Parametric Studies of Human Mars Entry, Descent, and Landing Systems

This paper will present a parametric analysis for entry, descent, and landing (EDL) concepts, enabling rapid systems assessment and tradespace exploration. The entry system uses a hypersonic inflatable aerodynamic decelerator (HIAD) technology. The baseline system includes elements for Mars aerocapture (AC) and EDL segments of the mission. The mission concept of operations (ConOps) begins at Mars arrival in a polar inclination. After performing an aerocapture (AC) maneuver into a pre-defined Mars parking orbit, the AC HIAD is jettisoned. The rest of thevehicle stays in the parking orbit for up to one year. The EDL sequence starts with a deorbit burn at the apoapsis of the parking orbit. After hypersonic entry with the EDL HIAD, the entry system uses a supersonic retropropulsion maneuver to slow the vehicle for the descent and landing segments of the mission. The vehicle will maintain a constant velocity of 2.5 m/s for 5 seconds prior to landing.The system includes a Mars Ascent Vehicle (MAV), Mars Descent Module (MDM), and two HIADs. The MDM includes a primary structure, tanks, engines, and radiators. The primary structure is an aluminum-lithium (Al-Li)cruciform design similar to the structural design of the Apollo Lunar Module. The cruciform planform layout results in four outer bays, with adequate volume in the corners between outer bays to package four landing gears. The central bay is reserved for packaging the MAV and the recessed MAV engines. Two of the outer bays accommodate main propellant tanks, with one LOX and one CH4 tank in each bay. The two remaining outer bays each house four rocket engine systems and associated support structure. Each HIAD comprises an inflatable structure, flexible thermal protection system, gas, and gas generators. The HIAD design used in this study is a stacked-toroid concept with pairing loop straps and radial/chevron straps. The baseline system lands a 20-t payload on the Mars surface. It is assumed the vehicle arrives at 6.2 km/s relative velocity at 90° inclination and is captured to a 1-Sol parking orbit.The baseline design includes many assumptions such as margins, arrival state, ConOps options, parking orbit, physical dimensions, propellant options, and technology concepts. The impact of these parameters are quantified through systems-level sensitivity analyses, which capture the global impact—not at a component level—but at the systems level. The systems-level sensitivities expose major design drivers and importance of each assumption for a design.Through tradespace exploration, a wide range of systems parameters are examined and compared for several feasible design options. Studies have been completed for the following input parameters: payload mass, propellant options, AC/EDL ballistic coefficient, lander thrust to weight ratio (T/W) (surrogate for the maximum EDL g’s),engine specific impulse (Isp), parking orbit, and inclination.The final paper will present and discuss the parametric approach used in the study. It will also include the results of recent systems analyses, sensitivity analyses, and tradespace exploration

Jamshid Samareh↗

NASA NTRS1992

The 1991-1992 project summaries

The Department of Mechanical Engineering at The University of Texas at Austin participated in seven cooperative design projects this year. Six of the projects were associated with the Johnson Space Center and include the design of a thermal control system for an inflatable lunar habitat module, a vibration isolation system for a space shuttle cycle ergometer, a radiator shading device for a lunar outpost, a reusable astronaut safety tether, a resistive exercise device for use on the space shuttle, and a fleet of autonomous regolith throwing devices for radiation shielding of lunar habitats. The seventh project is associated with the Jet Propulsion Lab and involves the design of a shock absorbing wheel for a small six-wheeled Martian rover vehicle.

Nichols, Steven P.↗

NASA NTRS2010

Space Medicine: Shuttle - Space Station Crew Health and Safety Challenges for Exploration

This slide presentation combines some views of the shuttle take off, and the shuttle and space station on orbit, and some views of the underwater astronaut training , with a general discussion of Space Medicine. It begins with a discussion of the some of the physiological issues of space flight. These include: Space Motion Sickness (SMS), Cardiovascular, Neurovestibular, Musculoskeletal, and Behavioral/Psycho-social. There is also discussion of the space environment and the issues that are posed including: Radiation, Toxic products and propellants, Habitability, Atmosphere, and Medical events. Included also is a discussion of the systems and crew training. There are also artists views of the Constellation vehicles, the planned lunar base, and extended lunar settlement. There are also slides showing the size of earth in perspective to the other planets, and the sun and the sun in perspective to other stars. There is also a discussion of the in-flight changes that occur in neural feedback that produces postural imbalance and loss of coordination after return.

Dervay, Joseph↗

NASA NTRS1991

A lunar rover powered by an orbiting laser diode array

A conceptual design of a high-power, long-duration lunar rover powered by a laser beam is proposed. The laser transmitter in lunar orbit consists of an SP-100 nuclear reactor prime power source providing 100 kW of electricity to a laser array that emits 50 kW of laser radiation. The laser radiation is beamed to the lunar surface where it is received by a GaAlAs solid-state, laser-to-electric converter. This converter provides 22 kW of electrical power to the rover vehicle for science, locomotion, and crew needs. The mass of the laser transmitter is approximately 5000 kg, whereas the mass of the rover power supply is 520 kg. The rover power unit is significantly less massive than alternative rover power units.

De Young, R. J.↗

NASA NTRS2007

Heliospheric Physics and NASA's Vision for Space Exploration

The Vision for Space Exploration outlines NASA's development of a new generation of human-rated launch vehicles to replace the Space Shuttle and an architecture for exploring the Moon and Mars. The system--developed by the Constellation Program--includes a near term (approx. 2014) capability to provide crew and cargo service to the International Space Station after the Shuttle is retired in 2010 and a human return to the Moon no later than 2020. Constellation vehicles and systems will necessarily be required to operate efficiently, safely, and reliably in the space plasma and radiation environments of low Earth orbit, the Earth's magnetosphere, interplanetary space, and on the lunar surface. This presentation will provide an overview of the characteristics of space radiation and plasma environments relevant to lunar programs including the trans-lunar injection and trans-Earth injection trajectories through the Earth's radiation belts, solar wind surface dose and plasma wake charging environments in near lunar space, energetic solar particle events, and galactic cosmic rays and discusses the design and operational environments being developed for lunar program requirements to assure that systems operate successfully in the space environment.

Minow, Joseph I.↗

NASA NTRSDate not supplied

Artemis IV Docking in Radiation Belt Charging Environment

NASA’s Artemis IV mission is planned to deliver the International Habitation Module (I-Hab) to the Gateway space station in lunar orbit. The I-Hab will be launched aboard the Space Launch System (SLS) vehicle as a co-manifested payload with the Orion spacecraft. After translunar injection begins, the Orion spacecraft will separate from the SLS Exploration Upper Stage (EUS) and then dock with the I-Hab to extract it from the EUS. Because of the altitude, orientation, and time for this to occur, the vehicle-to-vehicle potential between the I-Hab/EUS vehicle and Orion spacecraft could exceed several thousand volts. The docking of these two spacecrafts with such large differential potentials presents a challenge for the vehicles. This presentation describes the space environments at the docking altitude, the calculated vehicle-to-vehicle potential, and the possible impacts of the resulting voltage and current transients occurring at first contact. Additionally, possible risk mitigation tests to demonstrate compatibility with the transient current and voltages will be presented.

Matthew Mccollum↗

NASA NTRSDate not supplied

Laser Beam Welding Advancements for In-Space Servicing, Assembly, and Manufacturing

NASA is currently working to develop in-space servicing, assembly, and manufacturing (ISAM) capabilities for low Earth orbit and the lunar surface. One crucial technology for this effort is laser beam welding. Laser systems can perform joining, cleaning, cutting, and repair activities, which will enable the construction of large in-space structures that could not fit on a single launch vehicle, such as trusses for solar panels, radiators, or communications infrastructure. Multiple projects studying laser welding for space applications are currently underway at NASA Marshall Space Flight Center. One of these, the DIsk-Shaped Configurable and Modular vAcuum uNit (DISCMAN), is a compact vacuum chamber designed to support parameter development for laser welding in microgravity. It contains a rotating platen with weld samples made from aluminum, steel, and titanium, a high-power infrared laser, and integrated pumps for pulling vacuum inside the sample cartridge. The DISCMAN payload is planned to launch to the International Space Station, where welds will be performed under sustained microgravity inside the Bishop Airlock. Another effort underway at Marshall is the Lunar Assembly and Servicing by Autonomous Robotics (LASAR) initiative. This project uses a space-rated robotic arm equipped with a laser weld head, wire feeder, and multiple cameras to perform welds in a thermal vacuum chamber simulating the lunar surface environment. Some of these welds are done on snowflake joints, which are specially designed to slot together to join segments of trussN structures, allowing for the construction of tall surface infrastructure. DISCMAN, LASAR, and other projects are being carried out to advance the technological maturity of in-space laser beam welding, collect data to inform computational models, and learn reliable processes for creating weld joints in space. This work supports NASA’s greater goals to expand humanity’s presence in low Earth orbit, establish a permanent moon base, and eventually send crewed missions to Mars and beyond.

Manufacturing↗

NASA NTRSDate not supplied

Laser Beam Welding Advancements for In-Space Servicing, Assembly, and Manufacturing

Robotics↗

NASA NTRS1993

Thermal Protection System design studies for lunar crew module

The results of a study to predict aeroheating and Thermal Protection System (TPS) requirements for manned entry vehicles returning to Earth from the moon are presented. The effects of vehicle size and lunar-return strategies on the aerothermodynamic environment and TPS design were assessed. Study guidelines were based on an Apollo Command Module (CM) configuration and lunar return strategies included direct entry and aerocapture followed by Low Earth Orbit entry (LEO). Convective heating was obtained by the boundary layer integral matrix procedure (BLIMP) code, and radiative heating was computed with the QRAD program. The AESOP-STAB code and the AESOP-THERM code were used for TPS analysis for ablating materials and nonablating materials respectively. Results indicated that there was an optimum size for minimum heating and that direct entry had higher heating rates than aerocapture. Aerocapture resulted in higher heat loads and TPS weight. The TPS weight factor was 6-8 percent for all lunar return strategies, with the TPS weight being about 50 percent less than that of the original Apollo CM vehicle.

Williams, S. D.↗

NASA NTRSDate not supplied

Mars Opposition Piloted Nuclear Electric Propulsion (NEP)-Chem Vehicle

Many previous studies have examined sending crews to and from Mars. The most economical involved a ‘conjunction’ class whereby the crew spends around 500 days on Mars surface waiting for a ‘cheap’ return. The total mission time results in a mission duration around 3 years. Given the current demonstrated crew maximum of a 1 year stint on ISS, it is interesting to look at reducing that time to only two years, thus reducing risk and minimizing time in the Martian System. In order to meet such a short mission an ‘opposition’ class Mars mission (which includes a Venus flyby) was chosen. The energy required to perform such a mission in only two years (for the 2036 opportunity at least) is about three times that of the 3 year conjunction mission. The rocket equation clearly shows that this mission would then require several times the propellant of the three-year mission unless the Isp of the propulsion system can be increased. Electric propulsion can provide the 3-10x improvement in Isp but even with a nuclear reactor power levels could approach 10 MWe. As an alternative, a smaller reactor (1.5 MWe class) joined together with a chemical stage was found to allow for using each propulsion system to its best advantage: low thrust in interplanetary space and chemical in the gravity wells of Earth and Mars. Indeed, the use of high Isp, low thrust during the interplanetary leg of the journey’s reduced the required capture/departure ∆Vs by 5-10X. Lowering the NEP power also allowed fitting the power system into a single SLS launch – which limited the radiator area to ~ 2500m^2. For the first look a reactor using fuels created by the SP-100 program with a limit of ~1200K was assumed. Starting in the ‘Lunar Gateway’ also allowed for use of commercial tankers to fuel the vehicle in a relative benign place. A top level summary of the mission design, concept of operations, as well as a conceptual point design of the vehicle is described.

Nuclear Electric Propulsion↗

NASA NTRS1962

United States Space Explorations 1958

The film describes preparation and launch of five satellites and two space probes during 1958. On January 31, a Jupiter vehicle launched Explorer I into space. Data from this satellite was used to identify the van Allen radiation belts. On March 17, a Vanguard I rocket launched an Earth satellite with solar batteries. Data from the mission was used to determine that the Earth is slightly pear-shaped. On March 26, Explorer III was launched to further study the van Allen belts, micrometeoroid impacts, and internal and external temperatures. Explorer IV, launched on July 26, was intended to study radiation and temperature data. A lunar probe, ABLE I, was intended to measure radiation, magnetic fields of Earth and the Moon, density of micrometeoric matter, and internal temperatures. A four-stage rocket was used in the launch. However, a turbo-pump failed and the liquid oxygen pump stopped, resulting in a failed mission. On October 10, Pioneer I was launched by an ABLE vehicle. First and second stage velocity was less than desired and the probe did not leave Earth orbit. Attempts to attain escape velocity were unsuccessful. On December, a Jupiter boost vehicle was used to launch Juno II, with Pioneer III as the payload. Escape velocity was reached and Pioneer III left Earth's atmosphere. Failed launches, such as those of Vanguard boost vehicles and several Explorer satellites, also added to scientific knowledge.

Source record↗

NASA NTRSDate not supplied

Impact of Thermochromic Coatings on Thermal Management for Human Spacecraft Applications

Thermochromic variable emittance coatings (VECs) allow for passive, dynamic thermal management of space vehicles due to their temperature dependent optical properties. Ideally, a thermochromic material should exhibit low mid-IR emissivity for lower temperatures and switch to high mid-IR emissivity for higher temperatures; one promising material that demonstrates these characteristics is Lanthanum Strontium Manganite (LSM). The transition temperatures of LSM are dependent on the Lanthanum-to-Strontium ratio x, La1-xSrxMnO3. La0.8Sr0.2MnO3, which has an emittance of 0.5 for < 270 K, then a linear increase in emittance to 0.8 over an increase of 70 K where it will saturate at ~340K, is commonly utilized but this is tunable by changing the Sr concentration. The tunability of the transition temperature range of LSM motivates a need to understand the optimal thermal transition range for human spacecraft. Additionally, LSM has high solar absorptance and to combat this issue the VEC will be micropatterned on top of a highly solar reflective layer of Barium Sulfate (BaSO4). Consequently, this work will elucidate the tradeoffs between solar absorptance and variable emittance for this technology since the micropattern adversely couples the solar absorptance to the mid-IR emittance. This work develops a computational model using Thermal Desktop to study the utility of VECs on human spacecraft, such as Orion, subjected to mission phases such as lunar transit and Gateway docking. The model considers both external heat loads, from solar irradiation and planetary IR, as well as vehicle heat loads, from avionics and crew, as an input into the body mounted thermochromic radiators. There are several design targets for VECs, i.e. achieving the largest change in emittance, finding the optimal temperature range for the transition, and achieving the highest emittance in the high temperature phase. This work prioritizes these design targets for LSM to achieve the optimal heat rejection for a human spacecraft. The VECs are compared with static emittance coatings and the efficiency gains are quantified. Thermochromic VECs will ultimately lead to simplifications of active cooling systems which synergistically correlates to mass reduction and less energy consumption for space vehicles.

Joseph Peoples↗

NASA NTRSDate not supplied

Space Exploration and the Lunar Surface Environment

People and spacecraft exploring the Moon face unforgiving environments. The lunar surface is covered in fine, abrasive dust churned by meteoroid impacts. Explorers face extreme thermal, radiation, and lighting conditions. Landing spacecraft must account for erosion caused by rocket exhaust, which can be hazardous for the vehicle or nearby assets. This talk will survey these challenges and the work being done to overcome them to achieve mission success.

Wesley A. Chambers↗

NASA NTRS1990

Space vehicle propulsion systems - Environmental space hazards

An evaluation of hazards which exist in geo-lunar space and have the potential to negatively affect a long-term mission-oriented spacecraft systems is presented based on published data. The hazards are categorized as pervasive (radiation), incident specific (meteoroids and thermal shock), and chemically corrosive (monatomic oxygen). It appears that the number one priority should be the development of new materials; and the secondary concern should be the development of fabrication techniques for the exterior hull, so that incident specific hazards can be minimized in an active fashion. The pervasive hazard can be dealt with by exploring on-board circuit technology with ancillary monitoring systems. Effects of thermal shock on the exterior nozzle, directional gimbals, and internal combustion chamber geometry seem to need more investigation.

Bahr, G. K.↗

NASA NTRS1989

Preliminary design of a lunar construction utility vehicle

Construction of a lunar base, prior to manned occupancy, is one of the most demanding technological challenges facing space system designers today. A flexible lunar construction machine is needed that can be operated remotely and that can perform a variety of construction tasks over a wide range of lunar conditions. A preliminary lunar construction utility vehicle (LCUV) design has been developed as part of a capstone design course at Old Dominion University and is described in this summary report. The design requirements are taken from a 1988 USRA Summer Design Report entitled The Lunar Split Mission: A Robotic Constructed Lunar Base Scenario, and from the proceedings of a workshop hosted by United Technologies Corporation entitled Report of the In Situ Resources Utilization Workshop. The first report describes a bootstrap base concept in which a minimum of essential surface elements are delivered and configured such that minimum EVA is required to bring the initial base on-line. The base is to be built in three phases, the first of which will be unmanned, while the second and third will be manned. The key to these concepts is the development of a semiautonomous, telerobotic lunar construction utility vehicle (LCUV). The tasks required of this robotic vehicle during the phase 1 build-up are as follows: (1) surface element transportation, handling, and assembly; (2) soil excavation and movement for site preparation; (3) radiation protection and materials processing; and (4) repair and maintenance of surface elements. In order to meet the stated requirements, the LCUV must be: (1) transformable to perform a wide variety of tasks; (2) self supporting; (3) designed to allow for telerobotic control as well as autonomous operation; (4) able to transport one fully configured space station common module (SSCM); (5) upgradable to allow for future growth; and (6) easy to maintain.

Bryant, C. A.↗

NASA NTRS1972

Long range planning for the development of space flight emergency systems.

The importance of long-range planning for space flight emergency systems is pointed out. Factors in emergency systems planning are considered, giving attention to some of the mission classes which have to be taken into account. Examples of the hazards in space flight include fire, decompression, mechanical structure failures, radiation, collision, and meteoroid penetration. The criteria for rescue vehicles are examined together with aspects regarding the conduction of rescue missions. Future space flight programs are discussed, taking into consideration low earth orbit space stations, geosynchronous orbit space stations, lunar operations, manned planetary missions, future space flight vehicles, the space shuttle, special purpose space vehicles, and a reusable nuclear shuttle.

Bolger, P. H.↗

NASA NTRS2005

On Structural Design of a Mobile Lunar Habitat With Multi- Layered Environmental Shielding

This report presents an overview of a Mobile Lunar Habitat (MLH) structural design consisting of advanced composite materials. The habitat design is derived from the cylindrical-shaped U.S. Lab module aboard the International Space Station (ISS) and includes two lateral ports and a hatch at each end that geometrically match those of the ISS Nodes. Thus, several MLH units can be connected together to form a larger lunar outpost of various architectures. For enhanced mobility over the lunar terrain, the MLH uses six articulated insect-like robotic, retractable legs enabling the habitat to .t aboard a launch vehicle. The carbon-composite shell is sandwiched between two layers of hydrogen-rich polyethylene for enhanced radiation shielding. The pressure vessel is covered by modular double-wall panels for meteoroid impact shielding supported by externally mounted stiffeners. The habitat s structure is an assembly of multiple parts manufactured separately and bonded together. Based on the geometric complexity of a part and its material system, an appropriate fabrication process is proposed.

Pruitt, J. R.↗

NASA NTRS2026

The Operational NASA Anomaly Gas Analyzer and Exploration Follow-Ons

The NASA Anomaly Gas Analyzer (AGA) is the culmination of nearly ten years of advancement of core technology originally developed by Vista Photonics through the Small Business Innovation Research (SBIR) program and expanded using NASA program funding. The AGA is a portable, battery operated, optical gas detection instrument for continuous monitoring of H 2 O, CO 2 , O 2 , CO, NH 3 , HCN, HF, HCl in the spacecraft environment. Real time, in-situ monitored, gas concentrations are communicated at 1 Hz through a built-in display on-orbit and additionally through a serial interface on the ground. The instruments function over the typical range of temperatures and pressures required for the spacecraft environment. A ten-year operational life is targeted. Seven AGA instruments are currently deployed on the International Space Station and two Orion-specific units are manifested for Artemis 2. A variation of the operational AGA is under development for lunar Human Landing Systems. A wider calibrated operating pressure envelope is required in this application. Otherwise, the instruments straightforwardly address lessons learned from the original AGA by improving aspects of thermal control, ruggedness, and ease-of-use. A subset of AGA-developed sensed gases (H 2 O, CO 2 , O 2 ) is the subject of the Primary Constituent Monitor (PCM) development for the Habitation and Logistics Outpost (HALO) destined for lunar orbit. Two PCMs will be installed in HALO as part of the environmental control loop where they will interface with vehicle power and communications. While the basic sensor techniques are unchanged from the AGA, these instruments have been radiation hardened to survive for many years in the harsh target environment. Current development status for these instruments will be presented along with recent developments of a naval submarine variant.

Air Monitoring↗