SEARCH · Search NASA

Results for “Earth Entry System”

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.

SourceResearch connectionFrom yearThrough year

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 127 records · Page 7

All sources NASA NTRS DOE OSTI

NASA NTRS2010

Orion Entry Display Feeder and Interactions with the Entry Monitor System

The Orion spacecraft is designed to return astronauts to a landing within 10 km of the intended landing target from low Earth orbit, lunar direct-entry, and lunar skip-entry trajectories. Al pile the landing is nominally controlled autonomously, the crew can fly precision entries manually in the event of an anomaly. The onboard entry displays will be used by the crew to monitor and manually fly the entry, descent, and landing, while the Entry Monitor System (EMS) will be used to monitor the health and status of the onboard guidance and the trajectory. The entry displays are driven by the entry display feeder, part of the Entry Monitor System (EMS). The entry re-targeting module, also part of the EMS, provides all the data required to generate the capability footprint of the vehicle at any point in the trajectory, which is shown on the Primary Flight Display (PFD). It also provides caution and warning data and recommends the safest possible re-designated landing site when the nominal landing site is no longer within the capability of the vehicle. The PFD and the EMS allow the crew to manually fly an entry trajectory profile from entry interface until parachute deploy having the flexibility to manually steer the vehicle to a selected landing site that best satisfies the priorities of the crew. The entry display feeder provides data from the ENIS and other components of the GNC flight software to the displays at the proper rate and in the proper units. It also performs calculations that are specific to the entry displays and which are not made in any other component of the flight software. In some instances, it performs calculations identical to those performed by the onboard primary guidance algorithm to protect against a guidance system failure. These functions and the interactions between the entry display feeder and the other components of the EMS are described.

Baird, Darren↗

NASA NTRS1982

Atmospheric guidance techniques and performance

The Orbiter entry guidance system controls the Space Shuttle Orbiter from the initial atmospheric penetration point to the point at which an earth relative velocity of 2500 feet/second is reached. At the latter point, control of the Orbiter is transferred to the terminal area energy management system. The entry guidance system is based on the concept that the range to be flown during entry is a unique function of the drag deceleration profile flown throughout the entry. The range prediction during entry is based on analytic equations which are simple drag deceleration functions of earth relative velocity above Mach 10.5 and energy with respect to the earth below Mach 10.5. Flight through the entry corridor is accomplished by linking these simple drag deceleration functions together in series in order to define a drag deceleration reference profile. The results of the first three Space Shuttle missions have not only verified the entry guidance concept but have also demonstrated the stability of the guidance system.

Harpold, J. C.↗

NASA NTRS1997

Entry Dispersion Analysis for the Stardust Comet Sample Return Capsule

Stardust will be the first mission to return samples from beyond the Earth-Moon system. The sample return capsule, which is passively controlled during the fastest Earth entry ever, will land by parachute in Utah. The present study analyzes the entry, descent, and landing of the returning sample capsule. The effects of two aerodynamic instabilities are revealed (one in the high altitude free molecular regime and the other in the transonic/subsonic flow regime). These instabilities could lead to unacceptably large excursions in the angle-of-attack near peak heating and main parachute deployment, respectively. To reduce the excursions resulting from the high altitude instability, the entry spin rate of the capsule is increased. To stabilize the excursions from the transonic/subsonic instability, a drogue chute with deployment triggered by an accelerometer and timer is added prior to main parachute deployment. A Monte Carlo dispersion analysis of the modified entry (from which the impact of off-nominal conditions during the entry is ascertained) shows that the capsule attitude excursions near peak heating and drogue chute deployment are within Stardust program limits. Additionally, the size of the resulting 3-sigma landing ellipse is 83.5 km in downrange by 29.2 km in crossrange, which is within the Utah Test and Training Range boundaries.

Desai, Prasun N.↗

NASA NTRS1993

Anthropometric data from launch and entry suited test subjects for the design of a recumbent seating system

Returning space crews to Earth in a recumbent position requires the design of a new seating system. Current anthropometric data are based on measurements taken while the subjects were unsuited and sitting. To be most accurate, it is necessary to design by measurements of subjects in the launch and entry suit in a recumbent position. Since the design of the recumbent seating system must meet the requirements of both 5th percentile Japanese female and 95th percentile American male crew members, a delta is reported rather than absolute measurements of the test subjects. This delta is the difference in the measurements taken with the subjects unsuited and sitting and those taken with the subjects suited and recumbent. This delta, representative of the change due to the suit, can be added to the existing Man-Systems Integration Standards (NASA-STD-3000) anthropometric data to project the measurements for 5th percentile Japanese female and 95th percentile American male crew members. A delta accounting for the spinal elongation caused by prolonged exposures to microgravity is added as well. Both unpressurized and pressurized suit conditions are considered. Background information, the test protocol and procedure, analysis of the data, and recommendations are reported.

Stoycos, Lara E.↗

NASA NTRS2018

Exergy Efficiency of Interplanetary Transfer Vehicles

In order to optimize systems, systems engineers require some sort of measure with which to compare vastly different system components. One such measure is system exergy, or the usable system work. Exergy balance analysis models provide a comparison of different system configurations, allowing systems engineers to compare different systems configuration options. This paper presents the exergy efficiency of several Mars transportation system configurations, using data on the interplanetary trajectory, engine performance, and vehicle mass. The importance of the starting and final parking orbits is addressed in the analysis, as well as intermediate hyperbolic escape and entry orbits within Earth and Mars' spheres of influence (SOIs). Propulsion systems analyzed include low-enriched uranium (LEU) nuclear thermal propulsion (NTP), high-enriched uranium (HEU) NTP, LEU methane (CH4) NTP, and liquid oxygen (LOX)/liquid hydrogen (LH2) chemical propulsion.

Owen, Sean T.↗

NASA NTRS2018

Exergy Efficiency of Interplanetary Transfer Vehicles

Owen, Sean T.↗

NASA NTRS2018

Exergy Efficiency of Interplanetary Transfer Vehicles

Owen, Sean T.↗

NASA NTRS2018

ADEPT, A Mechanically Deployable Re-Entry Vehicle System, Enabling Interplanetary CubeSat and 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.

Adaptive Deployable Entry and Placement Technology↗

NASA NTRS2016

Passive vs. Parachute System Architecture for Robotic Sample Return Vehicles

The Multi-Mission Earth Entry Vehicle (MMEEV) is a flexible vehicle concept based on the Mars Sample Return (MSR) EEV design which can be used in the preliminary sample return mission study phase to parametrically investigate any trade space of interest to determine the best entry vehicle design approach for that particular mission concept. In addition to the trade space dimensions often considered (e.g. entry conditions, payload size and mass, vehicle size, etc.), the MMEEV trade space considers whether it might be more beneficial for the vehicle to utilize a parachute system during descent/landing or to be fully passive (i.e. not use a parachute). In order to evaluate this trade space dimension, a simplified parachute system model has been developed based on inputs such as vehicle size/mass, payload size/mass and landing requirements. This model works in conjunction with analytical approximations of a mission trade space dataset provided by the MMEEV System Analysis for Planetary EDL (M-SAPE) tool to help quantify the differences between an active (with parachute) and a passive (no parachute) vehicle concept.

Maddock, Robert W.↗

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 NTRS2013

Sabots, Obturator and Gas-In-Launch Tube Techniques for Heat Flux Models in Ballistic Ranges

For thermal protection system (heat shield) design for space vehicle entry into earth and other planetary atmospheres, it is essential to know the augmentation of the heat flux due to vehicle surface roughness. At the NASA Ames Hypervelocity Free Flight Aerodynamic Facility (HFFAF) ballistic range, a campaign of heat flux studies on rough models, using infrared camera techniques, has been initiated. Several phenomena can interfere with obtaining good heat flux data when using this measuring technique. These include leakage of the hot drive gas in the gun barrel through joints in the sabot (model carrier) to create spurious thermal imprints on the model forebody, deposition of sabot material on the model forebody, thereby changing the thermal properties of the model surface and unknown in-barrel heating of the model. This report presents developments in launch techniques to greatly reduce or eliminate these problems. The techniques include the use of obturator cups behind the launch package, enclosed versus open front sabot designs and the use of hydrogen gas in the launch tube. Attention also had to be paid to the problem of the obturator drafting behind the model and impacting the model. Of the techniques presented, the obturator cups and hydrogen in the launch tube were successful when properly implemented

obturators↗

NASA NTRS2020

Application of Risk Informed Decision Making to Highly Reliable Three Dimensionally Woven Thermal Protection System for Mars Sample Return

The NASA Risk Informed Decision Making process is used to assess a trade space of three dimensionally woven thermal protection systems for application to the Mars Sample Return Earth Entry Vehicle. Candidate architectures are assessed based on mission assurance, technical development, cost, and schedule risk. Assessment methodology differed between the architectures, utilizing a four-point quantitative scale for mission assurance and technical development and highly tailored PERT techniques for cost and schedule. Risk results are presented, in addition to a review of RIDM effectiveness for this application.

Needels, J.↗

NASA NTRS2020

Application of Risk Informed Decision Making to a Highly Reliable Three-Dimensionally Woven Thermal Protection System for Mars Sample Return

Needels, J.↗

NASA NTRS2014

Mars Sample Return Using Commercial Capabilities: Mission Architecture Overview

Mars Sample Return (MSR) is the highest priority science mission for the next decade as recommended by the recent Decadal Survey of Planetary Science. This paper presents an overview of a feasibility study for an MSR mission. The objective of the study was to determine whether emerging commercial capabilities can be used to reduce the number of mission systems and launches required to return the samples, with the goal of reducing mission cost. We report the feasibility of a complete and closed MSR mission design using the following scenario that covers three synodic launch opportunities, beginning with the 2022 opportunity: A Falcon Heavy injects a SpaceX Red Dragon capsule and trunk onto a Trans Mars Injection (TMI) trajectory. The capsule is modified to carry all the hardware needed to return samples collected on Mars including a Mars Ascent Vehicle (MAV), an Earth Return Vehicle (ERV), and hardware to transfer a sample collected in a previously landed rover mission to the ERV. The Red Dragon descends to land on the surface of Mars using Super Sonic Retro Propulsion (SSRP). After previously collected samples are transferred to the ERV, the single-stage MAV launches the ERV from the surface of Mars. The MAV uses a storable liquid bi-propellant propulsion system to deliver the ERV to a Mars phasing orbit. After a brief phasing period, the ERV, which also uses a storable bi-propellant system, performs a Trans Earth Injection (TEI) burn. Upon arrival at Earth, the ERV performs Earth and lunar swing-bys and is placed into a lunar trailing circular orbit - an Earth orbit, at lunar distance. A later mission, using Dragon and launched by a Falcon Heavy, performs a rendezvous with the ERV in the lunar trailing orbit, retrieves the sample container and breaks the chain of contact with Mars by transferring the sample into a sterile and secure container. With the sample contained, the retrieving spacecraft makes a controlled Earth re-entry preventing any unintended release of pristine martian materials into the Earth's biosphere. The analysis methods employed standard and specialized aerospace engineering tools. Mission system elements were analyzed with either direct techniques or by using parametric mass estimating relationships (MERs). The architecture was iterated until overall mission convergence was achieved on at least one path. Subsystems analyzed in this study include support structures, power system, nose fairing, thermal insulation, actuation devices, MAV exhaust venting, and GN&C. Best practice application of loads, mass growth contingencies, and resource margins were used. For Falcon Heavy capabilities and Dragon subsystems we utilized publically available data from SpaceX, published analyses from other sources, as well as our own engineering and aerodynamic estimates. Earth Launch mass is under 11 mt, which is within the estimated capability of a Falcon Heavy, with margin. Total entry masses between 7 and 10 mt were considered with closure occurring between 9 and 10 mt. Propellant mass fractions for each major phase of the EDL - Entry, Terminal Descent, and Hazard Avoidance - have been derived. An assessment of the effect of the entry conditions on the thermal protection system (TPS), currently in use for Dragon missions, shows no significant stressors. A useful payload mass of 2.0 mt is provided and includes mass growth allowances for the MAV, the ERV, and mission unique equipment. We also report options for the MAV and ERV, including propulsion systems, crewed versus robotic retrieval mission, as well as direct Earth entry. International planetary protection policies as well as verifiable means of compliance will have a large impact on any MSR mission design. We identify areas within our architecture where such impacts occur. We also describe preliminary compliance measures that will be the subject of future work. This work shows that emerging commercial capabilities as well as new methodologies can be used to efficiently support an important planetary science objective. The work also has applications for human exploration missions that use propulsive EDL techniques

Red Dragon↗

NASA NTRS2009

A Multidisciplinary Tool for Systems Analysis of Planetary Entry, Descent, and Landing (SAPE)

SAPE is a Python-based multidisciplinary analysis tool for systems analysis of planetary entry, descent, and landing (EDL) for Venus, Earth, Mars, Jupiter, Saturn, Uranus, Neptune, and Titan. The purpose of SAPE is to provide a variable-fidelity capability for conceptual and preliminary analysis within the same framework. SAPE includes the following analysis modules: geometry, trajectory, aerodynamics, aerothermal, thermal protection system, and structural sizing. SAPE uses the Python language-a platform-independent open-source software for integration and for the user interface. The development has relied heavily on the object-oriented programming capabilities that are available in Python. Modules are provided to interface with commercial and government off-the-shelf software components (e.g., thermal protection systems and finite-element analysis). SAPE runs on Microsoft Windows and Apple Mac OS X and has been partially tested on Linux.

Samareh, Jamshid A.↗

NASA NTRS2018

The HIAD Orbital Flight Demonstration 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. HIADs are 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 overcome the size constraints imposed on current rigid aeroshell entry systems. This permits use of larger aeroshells resulting in increased entry system performance (e.g. higher payload mass and/or volume, higher landing altitude at Mars). The Low Earth Orbit Flight Test of an Inflatable Decelerator (LOFTID) is currently scheduled for mid-2021. LOFTID will be launched out of Vandenberg Air Force Base as a secondary payload on an expendable launch vehicle. The flight test will employ a 6m diameter, 70-deg sphere-cone aeroshell and will provide invaluable high-energy orbital re-entry flight data. This data will be essential in supporting the HIAD team to mature the technology to diameters of 10m and greater. Aeroshells of this scale will address near-term commercial applications and potential future NASA missions.LOFTID will incorporate an extensive instrumentation suite totaling over 150 science measurements. This will include thermocouples, heat flux sensors, IR cameras, and a radiometer to characterize the aeroheating environment and aeroshell thermal response. An inertial measurement unit (IMU), GPS, and flush air data system will be included in order to reconstruct the flown trajectory and aerodynamic characteristics. Loadcells will be used to measure the HIAD structural loading, and HD cameras will be mounted on the aft segment looking at the aeroshell to monitor structural response. In addition to the primary instrumentation suite, a new fiber optic sensing system will be used to measure nose temperatures as a technology demonstration. The LOFTID instrumentation suites leverages Agency-wide expertise, with hardware development occurring at Ames Research Center, Langley Research Center, Marshall Space Flight Center and Armstrong Flight Research Center.This presentation will discuss the measurement objectives for the LOFTID mission, and the extensive instrumentation suite that has been selected to capture the HIAD's performance during the high-energy orbital re-entry flight test.

Swanson, Greg↗

NASA NTRS2018

The HIAD Orbital Flight Demonstration 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. HIADs are 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 overcome the size constraints imposed on current rigid aeroshell entry systems. This permits use of larger aeroshells resulting in increased entry system performance (e.g. higher payload mass and/or volume, higher landing altitude at Mars). The Low Earth Orbit Flight Test of an Inflatable Decelerator (LOFTID) is currently scheduled for mid-2021. LOFTID will be launched out of Vandenberg Air Force Base as a secondary payload on an expendable launch vehicle. The flight test will employ a 6m diameter, 70 degree sphere-cone aeroshell and will provide invaluable high-energy orbital re-entry flight data. This data will be essential in supporting the HIAD team to mature the technology to diameters of 10m and greater. Aeroshells of this scale will address near-term commercial applications and potential future NASA missions. LOFTID will incorporate an extensive instrumentation suite totaling over 150 science measurements. This will include thermocouples, heat flux sensors, IR cameras, and a radiometer to characterize the aeroheating environment and aeroshell thermal response. An inertial measurement unit (IMU), GPS, and flush air data system will be included in order to reconstruct the flown trajectory and aerodynamic characteristics. Loadcells will be used to measure the HIAD structural loading, and HD cameras will be mounted on the aft segment looking at the aeroshell to monitor structural response. In addition to the primary instrumentation suite, a new fiber optic sensing system will be used to measure nose temperatures as a technology demonstration. The LOFTID instrumentation suites leverages Agency-wide expertise, with hardware development occurring at Ames Research Center, Langley Research Center, Marshall Space Flight Center and Armstrong Flight Research Center. This presentation will discuss the measurement objectives for the LOFTID mission, and the extensive instrumentation suite that has been selected to capture the HIAD's performance during the high-energy orbital re-entry flight test.

Swanson, G. T.↗

NASA NTRS2003

Cardiovascular effects of anti-G suit and cooling garment during space shuttle re-entry and landing

BACKGROUND: Many cardiovascular changes associated with spaceflight reduce the ability of the cardiovascular system to oppose gravity on return to Earth, leaving astronauts susceptible to orthostatic hypotension during re-entry and landing. Consequently, an anti-G suit was developed to protect arterial pressure during re-entry. A liquid cooling garment (LCG) was then needed to alleviate the thermal stress resulting from use of the launch and entry suit. METHODS: We studied 34 astronauts on 22 flights (4-16 d). Subjects were studied 10 d before launch and on landing day. Preflight, crewmembers were suited with their anti-G suits set to the intended inflation for re-entry. Three consecutive measurements of heart rate and arterial pressure were obtained while seated and then again while standing. Three subjects who inflated the anti-G suits also donned the LCG for landing. Arterial pressure and heart rate were measured every 5 min during the de-orbit maneuver, through maximum G-loading (max-G) and touch down (TD). After TD, crew-members again initiated three seated measurements followed by three standing measurements. RESULTS: Astronauts with inflated anti-G suits had higher arterial pressure than those who did not have inflated anti-G suits during re-entry and landing (133.1 +/- 2.5/76.1 +/- 2.1 vs. 128.3 +/- 4.2/79.3 +/- 2.9, de-orbit; 157.3 +/- 4.5/102.1 +/- 3.6 vs. 145.2 +/- 10.5/95.7 + 5.5, max-G; 159.6 +/- 3.9/103.7 +/- 3.3 vs. 134.1 +/- 5.1/85.7 +/- 3.1, TD). In the group with inflated anti-G suits, those who also wore the LCG exhibited significantly lower heart rates than those who did not (75.7 +/- 11.5 vs. 86.5 +/- 6.2, de-orbit; 79.5 +/- 24.8 vs. 112.1 +/- 8.7, max-G; 84.7 +/- 8.0 vs. 110.5 +/- 7.9, TD). CONCLUSIONS: The anti-G suit is effective in supporting arterial pressure. The addition of the LCG lowers heart rate during re-entry.

NASA Center JSC↗