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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 307 records · Page 17

Are Samples Obtained after Return to Earth Reflective of Spaceflight or Increased Gravity?

Upon return to Earth, following space flight, living systems are immediately exposed to an increase in gravity of 1G. It has been difficult to differentiate between changes that are residuals of the acclimation to space flight from those resulting from acute exposure to an increase in =gravity upon re-entry. We compared previously reported changes observed in male Sprague-Dawley rats upon return to Earth to those induced by centrifugation, because both paradigms result in an increase of 1G. With both treatments there was a reduction in body mass, due to reduced food intake and increased urine output. The decrease in food intake was initially greater with centrifugation. The magnitudes of the changes in food intake and urine output were similar in both treatments. However, the slightly greater initial loss in body mass with centrifugation was due to a decrease in water intake not seen after space flight. The absence of pronounced differences between these treatments suggest the responses observed after landing are not residuals of adaptation to the space flight environment, but the result of adaptation to an increase in the level of gravity.

Wade, C. R.↗

Guidance and control analysis of the entry of a lifting body personnel launch vehicle

NASA is currently involved in definition studies of a Personnel Launch System (PLS) that could be used to transport people to and from low-earth orbit. This vehicle would serve both to complement the Space Shuttle and to provide alternative access to space in the event the Space Shuttle fleet were unavailable for a prolonged period. The PLS would consist of a manned spacecraft launched by an expendable vehicle, e.g., Titan 4. One promising candidate for the manned component of the PLS is the NASA Langley Research Center HL-20 lifting body. Many studies are currently underway to assess this vehicle, and one of the main areas of study is the development of the capability to successfully enter, glide to the landing site, and land. To provide this capability, guidance and control algorithms have been developed, incorporated into a six-degree-of-freedom simulation, and evaluation in the presence of off-nominal atmospheric conditions, consisting of both density variations and steady-state winds. In addition, the impact of atmospheric turbulence was examined for the portion of flight from Mach 3.5 to touchdown. This analysis showed that the vehicle remained controllable and could successfully land even in the presence of off-nominal atmospheric conditions.

Powell, Richard W.↗

ADEPT for Interplanetary Small Satellite Missions

There is growing interest for utilizing Small Satellites beyond low Earth orbit. A number of secondary CubeSat payload missions are planned at Mars, cis-Lunar Space, near Earth objects, and moons of the Gas Giants. Use of smaller systems may enable utilization of otherwise unused capacity of larger "host" missions. Development of re-entry systems that leverage and accommodate Small Satellite technology will substantially expand the range of mission applications by offering the capability for high speed entry or aerocapture at destinations with atmospheres. Deployable entry vehicles (DEVs) offer benefits over traditional rigid aeroshells including volume, mass and payload form factor. The Adaptive Deployable Entry and Placement Technology (ADEPT) offers such a delivery capability for Small Sat or CubeSat orbiter(s), in-situ elements, or landers. The ADEPT system can package with off the shelf CubeSat deployment systems (1U-16U) to offer a delivery capability for a single CubeSat or constellations. Furthermore, ADEPT can deliver the same science payload to a destination with a stowed diameter a factor of 3-4 times smaller than an equivalent rigid aeroshell, alleviating volumetric constraints on the secondary payload accommodation or primary carrier spacecraft bus. This paper will describe ADEPT's current development status and define various interplanetary mission concepts in order to provide guidelines for potential Small Satellite payload developers and mission implementers.

Cassell, Alan↗

OVERVIEW OF THE LOFTID INSTRUMENTATION SUITE

NASA’s Hypersonic Inflatable Aerodynamic Decelerator (HIAD) technology has been selected for a Technology Demonstration Mission under the Science and Technology Mission Directorate and is currently scheduled to launch in FY22. HIAD is an enabling technology that can facilitate atmospheric entry of heavy payloads to planets such as Earth and Mars using a deployable aeroshell. The deployable nature of the HIAD technology allows it to avoid the size constraints imposed on current rigid aeroshell entry systems. This enables use of larger aeroshells resulting in increased entry system performance (e.g. higher payload mass and/or volume, higher landing altitude at Mars).

G T Swanson↗

Cassini - A mission to the Saturnian system

This international (NASA/ESA) mission, planned for a Titan IV/Centaur G-prime launch in the 1995-1997 era, consists of a Saturn Orbiter (a U.S. Mariner Mark II spacecraft) and a Titan Atmospheric Entry Probe (an ESA product). The mission profile is highly sophisticated in its astrodynamic aspects - a deep space Delta-V maneuver, followed by earth and Jupiter gravity assist flybys, with asteroid encounter(s) along the way. Saturn arrival is characterized by further multiple maneuvers resulting in the Titan probe delivery and a close orbiter flyby, to be followed by a 4-year Saturn system tour. Details of the trajectory design process are discussed.

Sergeyevsky, Andrey B.↗

3D Woven Mid-Density Carbon Phenolic (3MDCP) Thermal Protection System Development

3-Dimensionally Woven, Mid-Density, Carbon Phenolic (3MDCP) Thermal Protection System (TPS) material is derived from the dual layer 3D woven Heatshield for Extreme Entry Environment Technology (HEEET) material. The baseline 3MDCP design is a single piece thermal protection system that avoids the manufacturing and certification challenges associated with a tiled configuration. 3MDCP is targeted for very aggressive entry environments such as high-speed sample return missions to Earth and missions to Saturn, Venus and the ice giants. A 3MDCP heatshield begins as a flat woven preform that is formed to a given heatshield shape and then infused with phenolic resin. NASA Ames, in collaboration with TEAM Inc. (weaving) and Fiber Materials Inc. (both Spirit AeroSystems Companies) have been developing and demonstrating the manufacturing processes to fabricate a 3MDCP heatshield at a diameter of 1.25 meters. The process of forming the flat woven preform into the final heatshield shape, a sphere-cone geometry, involves local movement of the yarns in the weave. This results in a single piece heatshield with continuous fibers, albeit with property variations between different regions on the heatshield. This presentation will provide a high-level status of 3MDCP development. This will include an overview of the manufacturing processes, with an emphasis on the impact of forming on fiber orientation, material properties and performance. The presentation will layout the plan for testing to assess the impact of forming on properties and review preliminary data comparing properties of flat to formed materials.

Thermal Protection System↗

Low-Earth Orbit Flight Test of an Inflatable Decelerator (LOFTID) Mission Overview, Science Return, and Future Applications of This Technology

The Low-Earth Orbit (LEO) Flight Test of an Inflatable Decelerator (LOFTID) mission was the culmination of two decades of research and development for Hyper-sonic Inflatable Aerodynamic Decelerator (HIAD) technology. LOFTID was a project overseen by the Technology Demonstration Mission (TDM) program within NASA’s Space Technology Mission Directorate. The success of the LOFTID mission could enable new NASA missions to Mars, Venus, and most solar-system destinations with atmospheres, as well as cost-effective payload returns to Earth, including in-space manufactured materials and launch vehicle asset recovery. LOFTID, with its unique inflatable heat shield, was the first-of-a-kind orbital reentry flight, and the largest blunt body atmospheric entry of any kind. On November 10, 2022, just over 10 years since the previous flight test of the smaller sub-orbital Inflatable Reentry Vehicle Experiment-3 (IRVE-3) [1], NASA Langley Research Center, with partner United Launch Alliance (ULA), successfully launched and achieved reentry and recovery of the LOFTID Reentry Vehicle (RV), further demonstrating the viability of the HIAD technology for large-diameter, inflatable heat shields to safely and accurately deliver large payloads through an atmosphere via a controlled descent and landing. Launching as a secondary payload with the Joint Polar Satellite System 2 (JPSS-2) from Vandenberg Space Force Base, California, stowed inside an ex-tended payload adapter of the Atlas V 401 launch vehicle, the LOFTID mission officially began after the JPSS-2 payload was delivered to its orbit by the Centaur second stage. The LOFTID RV was flying solo on its spin-stabilized ballistic reentry trajectory about one hour after launch, and the flight ended approximately one hour later with a gentle splashdown under parachute in the Pacific Ocean off the east coast of Hawaii, where the RV was recovered and later shipped back to NASA Langley. LOFTID endured the harsh environments of atmospheric reentry while demonstrating stable aerodynamics through the entire spectrum of hypersonic, supersonic, transonic, and subsonic flight. The LOFTID RV was exposed to an aeroheating environment representative of many Mars and LEO HIAD applications, while successfully demonstrating the ability of the heat-affected inflatable structure to with-stand aerodynamic forces that exceeded those expected at Mars. This flight demonstration of a 6m diameter HIAD confirmed the technology structural and thermal performance as it protected the 1100kg RV entering Earth’s atmosphere at 8 km/s, reaching Mach 30, and experiencing 9g deceleration before deploying parachutes and splashing down in the Pacific Ocean. HIAD technology involves an aeroshell that can be hard packed into a small volume for launch and then deployed prior to atmospheric entry for a controlled deceleration through the atmosphere. Large deployable heat shields enable spacecraft to carry bigger, heavier payloads, including scientific instruments and human support systems for planetary landing and exploration. Much larger than traditional fixed diameter aeroshells that are constrained by the size of launch vehicle shrouds, inflatable decelerators create more drag and start the deceleration process in the upper reaches of the atmosphere with greater efficiency and stability. LOFTID’s successful demonstration of the HIAD technology has greatly expanded NASA’s options for future planetary missions and opened new commercial opportunities for lower-cost mission pay-load recovery including recovery of launch vehicle assets, or as a delivery system to return cargo from the International Space Station or cislunar space. The developments for some of these applications are al-ready underway. While LOFTID was indeed a first-of-a-kind flight for an inflatable heatshield, the largest blunt-body entry ever, its remarkable performance assured that it will not be the last of its kind.

John DiNonno↗

Optimal trajectories for the aeroassisted flight experiment. Part 3: Formulation, results, and analysis

The determination of optimal trajectories for the aero-assisted flight experiment (AFE) is investigated. The intent of this experiment is to simulate a GEO-to-LEO transfer, where GEO denotes a geosynchronous Earth orbit and LEO denotes a low Earth orbit. The trajectories of an AFE spacecraft are analyzed in a 3D-space, employing the full system of 6 ODEs describing the atmospheric pass. The atmospheric entry conditions are given, and the atmospheric exit conditions are adjusted in such a way that the following conditions are satisfied: (1) the atmospheric velocity depletion is such that, after exiting, the AFE spacecraft first ascends to a specified apogee and then descends to a specified perigee; and (2) the exit orbital plane is identical with the entry orbital plane. The final maneuver, not analyzed here, includes the rendezvous with and the capture by the space shuttle.

Miele, A.↗

MMOD Protection and Degradation Effects for Thermal Control Systems

Micrometeoroid and orbital debris (MMOD) environment overview Hypervelocity impact effects & MMOD shielding MMOD risk assessment process Requirements & protection techniques - ISS - Shuttle - Orion/Commercial Crew Vehicles MMOD effects on spacecraft systems & improving MMOD protection - Radiators Coatings - Thermal protection system (TPS) for atmospheric entry vehicles Coatings - Windows - Solar arrays - Solar array masts - EVA Handrails - Thermal Blankets Orbital Debris provided by JSC & is the predominate threat in low Earth orbit - ORDEM 3.0 is latest model (released December 2013) - http://orbitaldebris.jsc.nasa.gov/ - Man-made objects in orbit about Earth impacting up to 16 km/s average 9-10 km/s for ISS orbit - High-density debris (steel) is major issue Meteoroid model provided by MSFC - MEM-R2 is latest release - http://www.nasa.gov/offices/meo/home/index.html - Natural particles in orbit about sun Mg-silicates, Ni-Fe, others - Meteoroid environment (MEM): 11-72 km/s Average 22-23 km/s.

Christiansen, Eric↗

Entry, Descent, and Landing Communications for the 2011 Mars Science Laboratory

The Mars Science Laboratory (MSL), established as the most advanced rover to land on the surface of Mars to date, launched on November 26th, 2011 and arrived to the Martian Gale Crater during the night of August 5th, 2012 (PDT). MSL will investigate whether the landing region was ever suitable to support carbon-based life, and examine rocks, soil, and the atmosphere with a sophisticated suite of tools. This paper addresses the flight system requirement by which the vehicle transmitted indications of the following events using both X-band tones and UHF telemetry to allow identification of probable root causes should a mission anomaly have occurred: Heat-Rejection System (HRS) venting, completion of the cruise stage separation, turn to entry attitude, atmospheric deceleration, bank angle reversal commanded, parachute deployment, heatshield separation, radar ground acquisition, powered descent initiation, rover separation from the descent stage, and rover release. During Entry, Descent, and Landing (EDL), the flight system transmitted a UHF telemetry stream adequate to determine the state of the spacecraft (including the presence of faults) at 8 kbps initiating from cruise stage separation through at least one minute after positive indication of rover release on the surface of Mars. The flight system also transmitted X-band semaphore tones from Entry to Landing plus one minute although since MSL was occulted, as predicted, by Mars as seen from the Earth, Direct-To-Earth (DTE) communications were interrupted at approximately is approx. 5 min after Entry ( approximately 130 prior to Landing). The primary data return paths were through the Deep Space Network (DSN) for DTE and the existing Mars network of orbiting assets for UHF, which included the Mars Reconnaissance Orbiter (MRO), Mars Odyssey (ODY), and Mars Express (MEX) elements. These orbiters recorded the telemetry data stream and returned it back to Earth via the DSN. The paper also discusses the total power received during EDL and the robustness of the telecom design strategy used to ensure EDL communications coverage.

relay operations↗

Mission and Design Sensitivities for Human Mars Landers Using Hypersonic Inflatable Aerodynamic Decelerators

Landing humans on Mars is one of NASA's long term goals. NASA's Evolvable Mars Campaign (EMC) is focused on evaluating architectural trade options to define the capabilities and elements needed to sustain human presence on the surface of Mars. The EMC study teams have considered a variety of in-space propulsion options and surface mission options. Understanding how these choices affect the performance of the lander will allow a balanced optimization of this complex system of systems problem. This paper presents the effects of mission and vehicle design options on lander mass and performance. Beginning with Earth launch, options include fairing size assumptions, co-manifesting elements with the lander, and Earth-Moon vicinity operations. Capturing into Mars orbit using either aerocapture or propulsive capture is assessed. For entry, descent, and landing both storable as well as oxygen and methane propellant combinations are considered, engine thrust level is assessed, and sensitivity to landed payload mass is presented. This paper focuses on lander designs using the Hypersonic Inflatable Aerodynamic Decelerators, one of several entry system technologies currently considered for human missions.

Polsgrove, Tara P.↗

High Ballistic Coefficient Mars EDL with Supersonic Retropropulsion

Future Mars missions will require the ability to land increasingly heavy systems with high precision, possibly at high altitude landing sites. A key obstacle in landing systems with high ballistic coefficient is Mars’ very thin atmosphere, approximately one hundred times less dense than Earth’s atmosphere. As a result, hypersonic deceleration occurs slowly throughout atmospheric entry and additional deceleration methods are required before landing. Previous landers up to and including MSL relied on Viking-heritage technologies including a Disk-Gap-Band (DGB) supersonic parachute and a blunt body aeroshell. Safe and effective deployment of a DGB parachute can be accomplished only within prescribed bounds on dynamic pressure and Mach number.2 In addition to the restrictive constraint this imposes on the entry trajectory, there is a concern that for a parachute to effectively slow an entry vehicle with a high ballistic coefficient, its diameter must be large enough to impose certification costs, modeling uncertainty, or mission risk that could be unacceptable. This has prompted investigation into alternative mission concepts based on use of supersonic retropropulsion (SRP) without a parachute. This work investigates SRP as an enabling technology for Mars EDL of high ballistic coefficient vehicles.

Noyes, Connor↗

Uncertainty Analysis of the Pitch Damping Coefficient of Blunt Bodies, Measured from Magnetic Suspension Wind Tunnel Tests

Blunt body aerodynamics in re-entry conditions are characterized by oscillatory behavior that is often encapsulated by the pitch damping coefficient (Cmq). Magnetic Suspension Wind Tunnel (MSWT) testing with a Magnetic Suspension Balance System (MSBS) aims to improve blunt body wind tunnel testing by removing the sometimes unknown aerodynamic effects of traditional sting mounts, as well as generate more constrained outputs than vertical spin tunnel testing. Efforts to calculate the uncertainty of the pitch damping coefficient with a proposed Earth Entry Vehicle (EEV) capsule model have been made by fitting an analytic prediction of the attitude time history to video data of a free-to-oscillate MSBS test.

Magnetic suspension, dynamic stability, uncertaint↗

An Overview of NASA’s Airborne and Field Data Resource Center

A key recommendation from NASA’s 2022 Airborne and Field Data Workshop called for the development of a virtual Resource Center for all stakeholders across the data lifecycle of airborne and field Earth observations. The agency’s Earth Science Data and Information System (ESDIS) Project and Airborne Data Management Group (ADMG) have worked in concert to establish the newly launched NASA Airborne and Field Data Resource Center (AFDRC) to provide a single entry point for a wide assortment of information on and the effective, responsible stewardship of non-satellite observational data. The AFDRC compiles access to many existing resources, but does so in a newly organized way that integrates availability to increase efficiency and holistic understanding while simplifying users’ experience. Initially launched in fall of 2023, the AFDRC is a NASA Earthdata domain website that clarifies several previously disparate resources and provides newly updated information, including: Learning Resources: Educational resources to broaden understanding of the role airborne and field observations play in advancing understanding of our planet and NASA’s role in collecting and archiving these data. Support for data users to Find and Access Data: Advanced contextual browse/search capabilities that efficiently link researchers to data products suitable for their science objectives - this includes linking to NASA’s Catalog of Archived Suborbital Earth Science Investigations (CASEI). Working with Data: Tools specific to individual types of suborbital Earth Science data and their (inter-)disciplinary communities to provide access as well as guidance for their application. Stewardship Responsibilities: Resources for data producers with to lessen requirement burdens at the time of data transfer, and information for data stewards with consistent, authoritative guidance on best practices and agency- and/or community- specific archival procedures. This presentation will give an overview of the motivation for NASA’s AFDRC, approach for the design and content, iterative community-driven improvements, promote the use of the AFDRC, and solicit additional feedback from airborne and field data user communities.

Sara Lubkin↗

Coronas-F Orbit Monitoring and Re-Entry Prediction

Russian scientific satellite CORONAS-F was launched on July, 31, 2001. The object was inserted in near-circular orbit with the inclination 82.5deg and a mean altitude approx. 520 km. Due to the upper atmosphere drag CORONAS-F was permanently descended and as a result on December, 6, 2005 it has finished the earth-orbital flight, having lifetime in space approx. 4.5 years. The satellite structural features and its flight attitude control led to the significant variations of its ballistic coefficient during the flight. It was a cause of some specific difficulties in the fulfillment of the ballistic and navigation support of this space vehicle flight. Besides the main mission objective CORONAS-F also has been selected by the Inter-Agency Space Debris Coordination Committee (IADC) as a target object for the next regular international re-entry test campaign on a program of surveillance and re-entry prediction for the hazard space objects within their de-orbiting phases. Spacecraft (S/C) CORONAS-F kept its working state right up to the end of the flight - down to the atmosphere entry. This fact enabled to realization of the additional research experiments, concerning with an estimation of the atmospheric density within the low earth orbits (LEO) of the artificial satellites, and made possible to continue track the S/C during final phase of its flight by means of Russian regular command & tracking system, used for it control. Thus there appeared a unique possibility of using for tracking S/C at its de-orbiting phase not only passive radar facilities, belonging to the space surveillance systems and traditionally used for support of the IADC re-entry test campaigns, but also more precise active trajectory radio-tracking facilities from the ground control complex (GCC) applied for this object. Under the corresponding decision of the Russian side such capability of additional high-precise tracking control of the CORONAS-F flight in this period of time has been implemented. The organizing of the CORONAS-F ballistic and navigational support (BNS) and solving its main tasks (such as S/C orbit determination (OD) and its motion prediction and connected with them) both for regular mission stage and for additional flight program were realized by the group of specialists from the Mission Control Center (MCC). MCC was also assigned as a principal organization from the Russian side for participation in the 7th IADC re-entry test campaign on CORONAS-F. The CORONAS-F flight features and space environments circumstances during its flight as well as a methodology and technology of spacecraft ballistic and navigational support are given below. The BNS results for different phases of S/C flight, including the results of its re-entry predictions, obtained during the realization of the 7th IADC test campaign are submitted. The accuracy of space vehicle re-entry prediction and its dependence on various factors are analyzed in more details.

Ivanov, N. M.↗

The Cassini/Huygens Mission to Saturn

The Cassini/Huygens mission is an international cooperative effort between NASA, the European Space Agency, and the Italian Space Agency to conduct a scientific investigation of the Saturnian system. The spacecraft, which was launched in October of 1997, will use a Venus-Venus-Earth-Jupiter gravity assisted trajectory to arrive at Saturn in July of 2004. The spacecraft, composed of an orbiter and an atmospheric entry probe, will be placed in orbit about Saturn, after which the Huygens probe will be released to enter the atmosphere of Titan, the largest moon of Saturn. The probe data is transmitted during its descent and possible survival on the surface to the orbiter overhead, where it is stored and later relayed to Earth. The orbiter then continues in a four-year mission about Saturn, conducting detailed studies of the atmosphere, the rings, the magnetosphere, Titan, and the icy satellites. The probe carries six instruments to explore the atmosphere and surface of Titan; the orbiter carries twelve instruments for its investigations. This paper characterizes the performance of the Cassini/Huygens spacecraft during its first year and three-quarters of flight, up through the second of the two Venus flybys. The flight system, including the 18 scientific instruments, is described, as well as the detailed overall objectives of the mission.

Mitchell, R.↗

Design, analysis, and tests of a shuttle-type heat-pipe-cooled leading edge

A one-half scale heat-pipe-cooled leading edge model was designed and fabricated to verify feasibility and performance of a full-scale Phase B shuttle orbiter design. Model performance was investigated experimentally by radiant heating and aerothermal tests and analytically by using a simple technique which predicts heat pipe start-up from the frozen state and also predicts transient and steady-state thermal behavior. Analytical results agree well with experimental results for start-up and steady-state heat pipe operation. The results indicate that variations in angle of attack and roll orientation had a negligible effect on heat pipe performance. The heat pipes effectively isothermalized the leading edge, and reduced peak temperatures to levels compatible with the use temperatures of superalloys. Results of these tests demonstrated the durability of the heat-pipe-cooled leading edge in withstanding earth-entry thermal and mechanical loads and indicate that the use of heat pipes in leading edge structures is a reliable concept for fully reusable hypersonic cruise and space transportation systems.

Camarda, C. J.↗

Comparison of mission design options for manned Mars missions

A number of manned Mars mission types, propulsion systems, and operational techniques are compared. Conjunction and opposition class missions for cryogenic, hybrid (cryo/storable), and NERVA propulsion concepts are addressed. In addition, both Earth and Mars orbit aerobraking, direct entry of landers, hyperbolic rendezvous, and electric propulsion cases are examined. A common payload to Mars was used for all cases. The basic figure of merit used was weight in low Earth orbit (LEO) at mission initiation. This is roughly proportional to launch costs.

Gus R Babb↗