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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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NASA NTRS1974

Approach guidance for outer planet Pioneer missions

An overall parametric system analysis of the use of an optical approach guidance measurement system (using a V-slit electrooptical 'star-pipper') onboard a Pioneer-type spin-stabilized spacecraft for outer planet missions. Optical measurements of satellites of Saturn and Uranus against the star background are considered. Measurement geometry and spacecraft attitude are evaluated in depth with respect to two Pioneer missions: Saturn-Uranus and Jupiter-Uranus atmospheric entry probe missions with departure from Earth in November/December, 1980. A preliminary evaluation of long-term attitude stability measurements of Pioneer 10 is presented in the Appendix. The major conclusions are: (1) at Saturn flyby, viewing some medium period satellites can provide the best measurement results. (2) At Uranus approach, viewing the planet may give best measurements if early bus-probe separation is required. (3) Final measurement accuracy depends on the data format transmitted to Earth and the applied data interpretation techniques.-

Bejczy, A. K.↗
NASA NTRS2020

Aerocapture as an Option for Ice Giants Mission

Aerocapture is an atmospheric maneuver where the aerodynamic forces of the vehicle (lift and drag) are used to provide the 𝞓V needed to slow down from the approach hyperbolic trajectory to achieve the desired captured orbit around the target planet. The aeroassist capture provides a large savings in propulsion needed to change the velocity of the vehicle, since aerodynamic forces rather than propulsive systems provide the change in velocity. Aerocapture requires an integrated system level design, including thermal protection systems, actuator systems for aerodynamic modulation, and guidance and control systems that can autonomously command the change in the aeroassist forces. Although aerocapture has been proposed for many situations in the past, including Mars, Venus, Titan, Saturn, Uranus, and Neptune orbiters as well as for Earth demonstration missions, it has not been attempted on any missions. However, many studies in the past three decades that have considered aerocapture as a design option have concluded that there are large mass savings that come from using aeroassist forces rather than propulsive forces to put a spacecraft into a captured orbit. The benefits are destination dependent, but some of the largest mass savings occur for the Ice Giants planets. Due to the large hyperbolic velocities of interplanetary trajectories approaching Uranus and Neptune, large amount of propulsion must be used to put a spacecraft in science orbits around these planets. Aerocapture can reduce the propulsion needs by dissipating energy in the sizable atmospheres of Uranus and Neptune. NASA commissioned a detailed study analysis to quantify the benefits, if any, for use at Neptune. The study found that even having to provide a heat shield aerocapture could deliver 40% more payload than an all-propulsive vehicle, and also provides for a 3–4-year reduction in trip time. Mass savings are expected also at Uranus with an aerocapture mission. Additionally, more recent advances in thermal protection systems and guidance and control systems show a path to increase capabilities beyond those results. This paper will discuss the merits of including aerocapture as an option for an Ice Giants mission. The discussion will focus on the cost and mass savings of using aerocapture instead of propulsive burns for Ice Giants orbiter mission scenarios while also detailing a potential concept of operations and entry vehicle design. Finally, the talk will discuss more recent work that shows aerocapture design is possible without the development of any novel entry vehicle. Thus, aerocapture can significantly improve science capabilities for an Ice Giants mission.

Soumyo Dutta↗
NASA NTRS1984

Mars exploration in the 1990's

Science objectives for a Mars exploration program in the 1990s, set forth by NASA's Committee on Planetary Exploration, are presented. They include: an intensive study of the Martian surface, the exploration of the structure and general circulation of the atmosphere, the exploration of the nature and dynamics of the planet's interior, the nature of Mars's magnetic field, the character of the upper atmosphere and its interaction with the solar wind, and the establishment of the global chemical and physical characteristics of the Martian surface. Several low cost missions, which have been suggested by the Committee for NASA's CORE program, such as the Mars Geoscience Climatology Observer, the Mars Aeronomy Orbiter, and the Mars Surface Network and Surface Probing missions are also covered. Direct Entry-Direct Return and Out-of-Orbit Entry-Orbital Rendezvous and Docking are two versions of a Mars Rover and Sample Return (MRSR) mission discussed in detail. It is concluded that a MRSR mission will require the combination of full Centaur capability, aerocapture at Mars and earth, and direct return to earth to meet major scientific goals. Further studies should be made in orbit assembly, fuel transfer technology, aeromaneuvering, terminal guidance, rover technology, and robotics.

Albee, A. L.↗
NASA NTRS1986

Electromagnetic braking for Mars spacecraft

Aerobraking concepts are being studied to improve performance and cost effectiveness of propulsion systems for Mars landers and Mars interplanetary spacecraft. Access to megawatt power levels (nuclear power coupled to high-storage inductive or capacitive devices) on a manned Mars interplanetary spacecraft may make feasible electromagnetic braking and lift modulation techniques which were previously impractical. Using pulsed microwave and magnetic field technology, potential plasmadynamic braking and hydromagnetic lift modulation techniques have been identified. Entry corridor modulation to reduce loads and heating, to reduce vertical descent rates, and to expand horizontal and lateral landing ranges are possible benefits. In-depth studies are needed to identify specific design concepts for feasibility assessments. Standing wave/plasma sheath interaction techniques appear to be promising. The techniques may require some tailoring of spacecraft external structures and materials. In addition, rapid response guidance and control systems may require the use of structurally embedded sensors coupled to expert systems or to artificial intelligence systems.

Holt, A. C.↗
NASA NTRS2001

2001 Flight Mechanics Symposium

This conference publication includes papers and abstracts presented at the Flight Mechanics Symposium held on June 19-21, 2001. Sponsored by the Guidance, Navigation and Control Center of Goddard Space Flight Center, this symposium featured technical papers on a wide range of issues related to attitude/orbit determination, prediction and control; attitude simulation; attitude sensor calibration; theoretical foundation of attitude computation; dynamics model improvements; autonomous navigation; constellation design and formation flying; estimation theory and computational techniques; Earth environment mission analysis and design; and, spacecraft re-entry mission design and operations. Government, industry, and the academic community participated in the preparation and presentation of these papers.

Flight mechanics↗
NASA NTRS2003

Atmospheric Models for Aerocapture

There are eight destinations in the Solar System with sufficient atmosphere for aerocapture to be a viable aeroassist option - Venus, Earth, Mars, Jupiter, Saturn and its moon Titan, Uranus, and Neptune. Engineering-level atmospheric models for four of these targets (Earth, Mars, Titan, and Neptune) have been developed for NASA to support systems analysis studies of potential future aerocapture missions. Development of a similar atmospheric model for Venus has recently commenced. An important capability of all of these models is their ability to simulate quasi-random density perturbations for Monte Carlo analyses in developing guidance, navigation and control algorithms, and for thermal systems design. Similarities and differences among these atmospheric models are presented, with emphasis on the recently developed Neptune model and on planned characteristics of the Venus model. Example applications for aerocapture are also presented and illustrated. Recent updates to the Titan atmospheric model, in anticipation of applications for trajectory and atmospheric reconstruct of Huygens Robe entry at Titan, are discussed. Recent updates to the Mars atmospheric model, in support of ongoing Mars aerocapture systems analysis studies, are also presented.

Justus, C. G.↗
NASA NTRS2003

Engineering-Level Model Atmospheres For Titan and Mars

An engineering-level atmospheric model for Titan has been developed for use in NASA s systems analysis studies of aerocapture and entry, descent and landing (EDL) applications in potential missions to Titan. Analogous to highly successful Global Reference Atmospheric Models for Earth (GRAM) and Mars (Mars-GRAM), the new model is called Titan-GRAM. Like GRAM and Mars-GRAM, an important feature of Titan-GRAM is its ability to simulate quasi-random perturbations for Monte-Carlo analyses in developing guidance, navigation and control algorithms, and for thermal systems design. Titan-GRAM features and sample results will be presented. Features of Mars-GRAM especially related to EDL applications will also be presented and illustrated.

Justus, C. G.↗
NASA NTRS2004

Atmospheric Models for Aerocapture

There are eight destinations in the solar System with sufficient atmosphere for aerocapture to be a viable aeroassist option - Venus, Earth, Mars, Jupiter, Saturn and its moon Titan, Uranus, and Neptune. Engineering-level atmospheric models for four of these targets (Earth, Mars, Titan, and Neptune) have been developed for NASA to support systems analysis studies of potential future aerocapture missions. Development of a similar atmospheric model for Venus has recently commenced. An important capability of all of these models is their ability to simulate quasi-random density perturbations for Monte Carlo analyses in developing guidance, navigation and control algorithm, and for thermal systems design. Similarities and differences among these atmospheric models are presented, with emphasis on the recently developed Neptune model and on planned characteristics of the Venus model. Example applications for aerocapture are also presented and illustrated. Recent updates to the Titan atmospheric model are discussed, in anticipation of applications for trajectory and atmospheric reconstruct of Huygens Probe entry at Titan.

Justus, C. G.↗
NASA NTRS2019

Aero Maneuvering Dynamics and Control for Precision Landing on Titan

We have developed and tested several dynamic models of the parafoil system descending in Titan’s atmosphere. These dynamic models include a progression from point mass models to rigid multibody models, including relative dynamics between canopy and payload. In these models, we have included wind models used for Titan simulation, and extrapolated wind gust models previously used for simulation in the Martian environment. We have also developed guidance and control techniques to for autonomous parafoil turning in the adverse wind environment. Finally, and in order to improve the controller performance by reducing the uncertainty to environmental factors, we have also developed ways to estimate the Titan environmental parameters, i.e. the atmospheric density, and the wind magnitude, during the descent. A more complete and realistic simulation is being developed, which uses JPL’s DSENDS Entry, Descent, and Landing Software framework.

Ermolli, Luca↗
NASA NTRS2021

Aerocapture Trajectories for Earth Orbit Technology Demonstration and Orbiter Science Missions at Venus, Earth, Mars, and Neptune

The use of aerocapture to provide the delta-V needed to capture a spacecraft into orbit can provide significant fuel savings compared to a propulsive orbit insertion. The aerocapture guidance method described in this paper uses drag modulation, which deploys a drag skirt in the atmosphere until the amount of delta-V needed to capture into the desired orbit is obtained, at which time the drag skirt is separated from the spacecraft. By modulating the time of drag skirt jettison, the vehicle can target a specific orbit apoapsis in the presence of uncertainties such as entry targeting errors and unknown atmospheric conditions.

Jadusingh, Matt↗
NASA NTRS2009

Astronaut Thermal Exposure: Re-Entry After Low Earth Orbit Rescue Mission

The STS-125 mission, launched May 11, 2009, is the final servicing mission to the Hubble Space Telescope. The repair mission's EVA tasks are described, including: installing a new wide field camera; installing the Cosmic Origins Spectrograph; repairing the Space Telescope Imaging Spectrograph; installing a new outer blanket layer; adding a Soft Capture and Rendezvous System for eventual controlled deorbit in about 2014; replacing the 'A' side Science Instrument Command and Data Handling module; repairing the Advanced Camera for surveys; and, replacing the rate sensor unit gyroscopes, fine guidance sensors and 3 batteries. Additionally, the Shuttle crew cabin thermal environment is described. A CFD model of per person CO2 demonstrates a discrepancy between crew breathing volume and general mid-deck levels of CO2. A follow-on CFD analysis of the mid-deck temperature distribution is provided. Procedural and engineering mitigation plans are presented to counteract thermal exposure upon reentry to the Earth atmosphere. Some of the procedures include: full cold soak the night prior to deorbit; modifying deck stowage to reduce interference with air flow; and early securing of avionics post-landing to reduce cabin thermal load prior to hatch opening. Engineering mitigation activities include modifying the location of the aft starboard ICUs, eliminating the X3 stack and eliminating ICU exhaust air directed onto astronauts; improved engineering data of ICU performance; and, verifying the adequacy of mid-deck temperature control using CFD models in addition to lumped parameter models. Post-mitigation CFD models of mid-deck temperature profiles and distribution are provided.

Gillis, David B.↗
NASA NTRSDate not supplied

Aerocapture Trajectory Design for Uranus Orbiter

Introduction: The recently released National Academies Planetary Science and Astrobiology Decadal Survey 2023-2032 [1] identified the Ice Giants as the top priority science destination. While the survey acknowledged the potential for either a Uranus Orbiter and Probe (UOP) mission or a Neptune-Triton Odyssey mission, it ultimately identified the former as the highest priority new flagship mission. UOP missions calls for a launch window of opportunity between 2031-2038 with 12-15 year interplanetary cruise time along with a fully-propulsive Uranus Orbit Insertion burn on the order of a few km/s. However, a mission to Uranus with the same science payload could utilize aerocapture for orbit insertion to achieve both a significant reduction in the interplanetary cruise time and reduction in propulsive burn costs. Why Aerocapture: Aerocapture is a promising propellant and time-saving orbital insertion technique for planetary destinations with an atmosphere. Although not flight-proven, previous aerocapture systems studies in the literature have demonstrated both the validity and robustness of the technique at various planetary destinations. With respect to the Ice Giant planets, Neptune has seen more of the analysis in the literature. For science missions at Neptune, aerocapture can enable 1.4 times more delivered mass to orbit than an all-propulsive mission for the same launch vehicle while reducing interplanetary cruise times by more than 3 years [2] Additionally with modern guidance and control, Neptune aerocapture with blunt-body aeroshells is realizable [3][4]. There are limited papers in the literature investigating Uranus aerocapture with those available providing a preliminary feasibility assessment [5]. Consequently, the two-year funded NASA Space Technology Mission Directorate (STMD)-funded project, titled Aerocapture System as an Enabling Technology for Ice Giants Missions, aims to mature the analysis and technology state of Uranus aerocapture. Trajectory Design: This paper presents the current state of the trajectory design in support of the new aerocapture project. The project design philosophy is inspired from recent Neptune aerocapture studies, which employed modern guidance and control, in the sense that blunt-body aeroshells are analyzed. An assessment of the theoretical flight path angle corridor width is conducted for a range of ballistic coefficients and lift-to-drag ratios for both Space Launch System and Falcon Heavy Launch Vehicle interplanetary trajectory solutions. The results from the corridor width assessment provide an assessment of the aerocapture design-space and qualitative metrics on trajectory design considerations. The Program to Optimize Simulated Trajectories II (POST2) is utilized to run Monte Carlo simulations of Uranus aerocapture three-degree-of-freedom bank angle modulated trajectories using a closed-loop numerical-predictor corrector guidance algorithm. UranusGRAM 2021 is utilized as the atmospheric model [6]. A Uranus-developed aerodatabase, originally derived from Mars Science Laboratory (MSL), is utilized to provide vehicle aerodynamics over a wide range of hypersonic flow regimes. A MSL-derived 70 deg 4.5m diameter sphere-cone aeroshell that houses the UOP payload mass is assumed. Robustness testing and performance analysis is conducted, including the assessment of entry state errors, atmosphere density variations, and aerodynamic dispersions. Post-aerocapture Delta-V and aerothermal statistics are formulated into propellant mass and TPS requirements. The results presented in the paper will demonstrate the trajectory viability of Uranus aerocapture. Preliminary Results: Preliminary trajectory design results indicates successful Uranus aerocapture with a blunt-body aeroshell housing the same payload mass as the UOP mission from an Earth-to-Uranus interplanetary trajectory arriving in less than 7 years. From this interplanetary trajectory, aerocapture provides an orbit insertion Delta-V capability of 6.9 km/s requiring less than 300 m/s for post-aerocapture correction burns (15% of wet mass allocated to propellant). To put this into perspective, the UOP study utilized an Earth-to-Uranus trajectory that arrives in 13 years and requires more than 1000 m/s for fully-propulsive orbit insertion (40% of wet mass allocated to propellant). Achieving the same 6.9 km/s Delta-V capability fully-propulsively is mass prohibitive (97% of wet mass allocated to propellant). Nevertheless, aerocapture has the potential to reduce interplanetary transit times to Uranus by half while delivering the same payload mass to orbit in a reduced propellant mass footprint.

Rohan Deshmukh↗
NASA NTRS2000

Simulators for the X-38/CRV Parafoil and Re-Entry Phases

The X38/CRV is unique in several aspects: it does not afford the pilot a forward view through a window and it utilizes a parafoil in the final landing flight phase. As a result of some of these unique attributes the prototype on-board displays will need to provide enhanced situation awareness to aid a crew member in supervising the onboard GNC software to select a landing site, and possible obstacle avoidance during the final flight phase. In the ESA/ESTEC visualization prototype system the scene is visualized using MultiGen OpenFlight models on high-end SGI machines. The Parafoil Guidance Navigation and Control (PGNC) algorithms are being developed as a joint NASA/ESA venture. These algorithms will be used onboard the vehicle to create nominal flight trajectories for the vehicle to follow. These same algorithms are also used as parafoil simulators in our augmented situation awareness system we are developing. The output of the parafoil simulator is displayed on a Windows-based computer running the LandForm FlightVision software. The FlightVision software is used to create the synthetic environment displays and all the necessary HUD symbology. Maps, such as aeronautical charts, as well as satellite imagery are optionally overlaid on the 3D-terrain model to provide additional situation awareness for the crew.

SITUATIONAL AWARENESS↗
NASA NTRS1976

Preliminary input to the space shuttle reaction control subsystem failure detection and identification software requirements (uncontrolled)

The current baseline method and software implementation of the space shuttle reaction control subsystem failure detection and identification (RCS FDI) system is presented. This algorithm is recommended for conclusion in the redundancy management (RM) module of the space shuttle guidance, navigation, and control system. Supporting software is presented, and recommended for inclusion in the system management (SM) and display and control (D&C) systems. RCS FDI uses data from sensors in the jets, in the manifold isolation valves, and in the RCS fuel and oxidizer storage tanks. A list of jet failures and fuel imbalance warnings is generated for use by the jet selection algorithm of the on-orbit and entry flight control systems, and to inform the crew and ground controllers of RCS failure status. Manifold isolation valve close commands are generated in the event of failed on or leaking jets to prevent loss of large quantities of RCS fuel.

Bergmann, E.↗
NASA NTRS1990

Decomposition technique and optimal trajectories for the aeroassisted flight experiment

An actual geosynchronous Earth orbit-to-low Earth orbit (GEO-to-LEO) transfer is considered with reference to the aeroassisted flight experiment (AFE) spacecraft, and optimal trajectories are determined by minimizing the total characteristic velocity. The optimization is performed with respect to the time history of the controls (angle of attack and angle of bank), the entry path inclination and the flight time being free. Two transfer maneuvers are considered: direct ascent (DA) to LEO and indirect ascent (IA) to LEO via parking Earth orbit (PEO). By taking into account certain assumptions, the complete system can be decoupled into two subsystems: one describing the longitudinal motion and one describing the lateral motion. The angle of attack history, the entry path inclination, and the flight time are determined via the longitudinal motion subsystem. In this subsystem, the difference between the instantaneous bank angle and a constant bank angle is minimized in the least square sense subject to the specified orbital inclination requirement. Both the angles of attack and the angle of bank are shown to be constant. This result has considerable importance in the design of nominal trajectories to be used in the guidance of AFE and aeroassisted orbital transfer (AOT) vehicles.

Miele, A.↗
NASA NTRS1991

Decomposition technique and optimal trajectories for the aeroassisted flight experiment

An actual geosynchronous earth orbit-to-low earth orbit (GEO-to-LEO) transfer is considered with reference to the aeroassisted flight experiment (AFE) spacecraft, and optimal trajectories are determined by minimizing the total characteristic velocity. The optimization is performed with respect to the time history of the controls (angle of attack and angle of bank), the entry path inclination and the flight time being free. Two transfer maneuvers are considered: direct ascent (DA) to LEO and indirect ascent (IA) to LEO via parking earth orbit (PEO). By taking into account certain assumptions, the complete system can be decoupled into two subsystems: one describing the longitudinal motion and one describing the lateral motion. The angle of attack history, the entry path inclination, and the flight time are determined via the longitudinal motion subsystem. In this subsystem, the difference between the instantaneous bank angle and a constant bank angle is minimized in the least square sense subject to the specified orbital inclination requirement. Both the angles of attack and the angle of bank are shown to be constant. This result has considerable importance in the design of nominal trajectories to be used in the guidance of AFE and aeroassisted orbital transfer (AOT) vehicles.

Miele, A.↗
NASA NTRS2011

Assessment of Laminar, Convective Aeroheating Prediction Uncertainties for Mars Entry Vehicles

An assessment of computational uncertainties is presented for numerical methods used by NASA to predict laminar, convective aeroheating environments for Mars entry vehicles. A survey was conducted of existing experimental heat-transfer and shock-shape data for high enthalpy, reacting-gas CO2 flows and five relevant test series were selected for comparison to predictions. Solutions were generated at the experimental test conditions using NASA state-of-the-art computational tools and compared to these data. The comparisons were evaluated to establish predictive uncertainties as a function of total enthalpy and to provide guidance for future experimental testing requirements to help lower these uncertainties.

Hollis, Brian R.↗
NASA NTRS2013

Assessment of Laminar, Convective Aeroheating Prediction Uncertainties for Mars-Entry Vehicles

An assessment of computational uncertainties is presented for numerical methods used by NASA to predict laminar, convective aeroheating environments for Mars-entry vehicles. A survey was conducted of existing experimental heat transfer and shock-shape data for high-enthalpy reacting-gas CO2 flows, and five relevant test series were selected for comparison with predictions. Solutions were generated at the experimental test conditions using NASA state-of-the-art computational tools and compared with these data. The comparisons were evaluated to establish predictive uncertainties as a function of total enthalpy and to provide guidance for future experimental testing requirements to help lower these uncertainties.

Hollis, Brian R.↗