Engineering Papers⌕ Search

SEARCH · Engineering Papers

Results for “Entry Performance”

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.

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

At least 37 records · Page 2
All sourcesNASA NTRSDOE OSTI
NASA NTRS2012

Post-Flight EDL Entry Guidance Performance of the 2011 Mars Science Laboratory Mission

The 2011 Mars Science Laboratory was the first successful Mars mission to attempt a guided entry which safely delivered the rover to a final position approximately 2 km from its target within a touchdown ellipse of 19.1 km x 6.9 km. The Entry Terminal Point Controller guidance algorithm is derived from the final phase Apollo Command Module guidance and, like Apollo, modulates the bank angle to control the range flown. For application to Mars landers which must make use of the tenuous Martian atmosphere, it is critical to balance the lift of the vehicle to minimize the range error while still ensuring a safe deploy altitude. An overview of the process to generate optimized guidance settings is presented, discussing improvements made over the last nine years. Key dispersions driving deploy ellipse and altitude performance are identified. Performance sensitivities including attitude initialization error and the velocity of transition from range control to heading alignment are presented. Just prior to the entry and landing of MSL in August 2012, the EDL team examined minute tuning of the reference trajectory for the selected landing site, analyzed whether adjustment of bank reversal deadbands were necessary, the heading alignment velocity trigger was in union with other parameters to balance the EDL risks, and the vertical L/D command limits. This paper details a preliminary postflight assessment of the telemetry and trajectory reconstruction that is being performed, and updates the information presented in the former paper Entry Guidance for the 2011 Mars Science Laboratory Mission (AIAA Atmospheric Flight Mechanics Conference; 8-11 Aug. 2011; Portland, OR; United States)

Mendeck, Gavin F.↗
NASA NTRS1980

Performance of entry heat shields on Pioneer Venus probes

The Pioneer Venus probes approached Venus with high relative velocity. As they entered the atmosphere, they were rapidly decelerated by aerodynamic drag, and a great deal of heat was generated. To protect the probe structure and the scientific instruments, a carbon phenolic heat shield was placed on the front of the probes. Because the design of heat shields for planetary entry is a developing technology, thermocouples were placed in the heat shields so that actual and predicted heat shield performance could be compared. The function of the heat shield is discussed, the probe environments during entry into the Venusian atmosphere are described, and some results from the heat shield experiment are presented. It was found that for the most part, the heat shields performed better than expected.

Pitts, W. C.↗
NASA NTRS1973

Balloon launched Viking decelerator test program

Four BLDT flights were conducted during the summer of 1972. The purpose of these tests was to qualify the Viking parachute system behind the full-scale Viking entry vehicle over the maximum range of entry conditions anticipated in the Viking '75 soft landing on Mars. A summary of the test series is presented. Test conditions ranged from a Mach number of 2.0 to 0.5 and dynamic pressure from 11.7 to 4.4 psf. This range of conditions covers the uncertainty in entry conditions at Mars due to atmospheric and entry performance uncertainties. Emphasis is placed on parachute performance and simulated Mars entry vehicle motions as influenced by the parachute performance. Conclusions are presented regarding the ability of the parachute to perform within the operational parameters required for a successful soft Martian landing. A list of references which covers all reports in the qualification test program is included.

Moog, R. D.↗
NASA NTRS2013

Post-Flight EDL Entry Guidance Performance for the 2011 Mars Science Laboratory Mission

Mars Science Laboratory landed Curiosity in Gale Crater on 5 Aug 2012. Curiosity landed successfully only 2.2 km away from the expected target given the onboard navigation state. Better than the average Apollo capsule splashdown miss distance. A late bank reversal and a suspected tail wind contributed to this slight miss. Entry guidance is derived from the Apollo capsule "final phase" logic and adjusts the range flown during entry by varying the direction of the lift vector (i.e., bank angle). Refinement of the guidance gains and alternative parachute deploy triggers to reduce the ellipse size will be studied for future Mars landing missions.

Mendeck, G.↗
NASA NTRSDate not supplied

Comparative Performance of Entry Simulations using Guided Human-Scale Landers at Mars

There has been ongoing interest in human exploration of Mars since the 1970s. Most studies to date have investigated moderate lift-to-drag ratio vehicles to investigate this problem. This paper considers the development and cross-simulation comparison of a slender-bodied, moderate lift-to-drag human-scale lander vehicle for Mars. The vehicle employs the Fully Numerical Predictor-corrector Entry Guidance to target a powered descent state, subsequently leveraging Fractional Polynomial Powered Descent Guidance laws to reach the desired touchdown target. Differences in simulation environment as well as nominal trajectory results are shared.

Evan Roelke↗
NASA NTRSDate not supplied

Comparative Performance of Entry Simulations using Guided Human-Scale Landers at Mars

There has been ongoing interest in human exploration of Mars since the 1970s. Most studies to date have investigated moderate lift-to-drag ratio vehicles to investigate this problem. This paper considers the development and cross-simulation comparison of a slender-bodied, moderate lift-to-drag human-scale lander vehicle for Mars. The vehicle employs the Fully Numerical Predictor-corrector Entry Guidance to target a powered descent state, subsequently leveraging Fractional Polynomial Powered Descent Guidance laws to reach the desired touchdown target. Differences in simulation environment as well as nominal trajectory results are shared.

Evan Roelke↗
NASA NTRS2005

An Approximate Ablative Thermal Protection System Sizing Tool for Entry System Design

A computer tool to perform entry vehicle ablative thermal protection systems sizing has been developed. Two options for calculating the thermal response are incorporated into the tool. One, an industry-standard, high-fidelity ablation and thermal response program was integrated into the tool, making use of simulated trajectory data to calculate its boundary conditions at the ablating surface. Second, an approximate method that uses heat of ablation data to estimate heat shield recession during entry has been coupled to a one-dimensional finite-difference calculation that calculates the in-depth thermal response. The in-depth solution accounts for material decomposition, but does not account for pyrolysis gas energy absorption through the material. Engineering correlations are used to estimate stagnation point convective and radiative heating as a function of time. The sizing tool calculates recovery enthalpy, wall enthalpy, surface pressure, and heat transfer coefficient. Verification of this tool is performed by comparison to past thermal protection system sizings for the Mars Pathfinder and Stardust entry systems and calculations are performed for an Apollo capsule entering the atmosphere at lunar and Mars return speeds.

Dec, John A.↗
NASA NTRS2006

An Approximate Ablative Thermal Protection System Sizing Tool for Entry System Design

A computer tool to perform entry vehicle ablative thermal protection systems sizing has been developed. Two options for calculating the thermal response are incorporated into the tool. One, an industry-standard, high-fidelity ablation and thermal response program was integrated into the tool, making use of simulated trajectory data to calculate its boundary conditions at the ablating surface. Second, an approximate method that uses heat of ablation data to estimate heat shield recession during entry has been coupled to a one-dimensional finite-difference calculation that calculates the in-depth thermal response. The in-depth solution accounts for material decomposition, but does not account for pyrolysis gas energy absorption through the material. Engineering correlations are used to estimate stagnation point convective and radiative heating as a function of time. The sizing tool calculates recovery enthalpy, wall enthalpy, surface pressure, and heat transfer coefficient. Verification of this tool is performed by comparison to past thermal protection system sizings for the Mars Pathfinder and Stardust entry systems and calculations are performed for an Apollo capsule entering the atmosphere at lunar and Mars return speeds.

Dec, John A.↗
NASA NTRSDate not supplied

Guidance Performance of Bank-Angle and Alpha-Beta Steering for A Large Robotic Mars Entry Vehicle

Entry performance using Apollo-based entry guidance is assessed with a future entry system for a large robotic payload. Both a traditional bank-angle steering system and an innovative alpha-beta steering concept are considered to compare flight performance. Results indicate alpha-beta steering provides an order of magnitude improvement in crossrange error and mean altitudes up to 1 km higher than bank-angle steering. Results also show alpha-beta steering provides downrange errors with a spread half of that of bank-angle steering. Lastly, alpha-beta steering is found to provide a narrower spread of terminal altitude, Mach number, and dynamic pressure, indicating safer parachute deploy conditions.

Daniel L Engel↗
NASA NTRS2009

Aerodynamic Challenges for the Mars Science Laboratory Entry, Descent and Landing

An overview of several important aerodynamics challenges new to the Mars Science Laboratory (MSL) entry vehicle are presented. The MSL entry capsule is a 70 degree sphere cone-based on the original Mars Viking entry capsule. Due to payload and landing accuracy requirements, MSL will be flying at the highest lift-to-drag ratio of any capsule sent to Mars (L/D = 0.24). The capsule will also be flying a guided entry, performing bank maneuvers, a first for Mars entry. The system's mechanical design and increased performance requirements require an expansion of the MSL flight envelope beyond those of historical missions. In certain areas, the experience gained by Viking and other recent Mars missions can no longer be claimed as heritage information. New analysis and testing is re1quired to ensure the safe flight of the MSL entry vehicle. The challenge topics include: hypersonic gas chemistry and laminar-versus-turbulent flow effects on trim angle, a general risk assessment of flying at greater angles-of-attack than Viking, quantifying the aerodynamic interactions induced by a new reaction control system and a risk assessment of recontact of a series of masses jettisoned prior to parachute deploy. An overview of the analysis and tests being conducted to understand and reduce risk in each of these areas is presented. The need for proper modeling and implementation of uncertainties for use in trajectory simulation has resulted in a revision of prior models and additional analysis for the MSL entry vehicle. The six degree-of-freedom uncertainty model and new analysis to quantify roll torque dispersions are presented.

Schoenenberger, Mark↗
NASA NTRS2007

NASA CEV Reference Entry GN&C System and Analysis

As part of its overall objectives, the Orion spacecraft will be required to perform entry and Earth landing functions for Low Earth Orbit (LEO) and Lunar missions. Both of these entry scenarios will begin with separation of the Service Module (SM), making them unique from other Orion mission phases in that only the Command Module (CM) portion of the Crew Exploration Vehicle (CEV) will be involved, requiring a CM specific Guidance, Navigation and Control (GN&C) system. Also common to these mission scenarios will be the need for GN&C to safely return crew (or cargo) to earth within the dynamic thermal and structural constraints of entry and within acceptable accelerations on the crew, utilizing the limited aerodynamic performance of the CM capsule. The lunar return mission could additionally require an initial atmospheric entry designed to support a precision skip and second entry, all to maximize downrange performance and ensure landing in the United States. This paper describes the Entry GN&C reference design, developed by the NASA-led team, that supports these entry scenarios and that was used to validate the Orion System requirements. Description of the reference design will include an overview of the GN&C functions, avionics, and effectors and will relate these to the specific design drivers of the entry scenarios, as well as the desire for commonality in vehicle systems to support the different missions. The discussion will also include the requirement for an Emergency Entry capability beyond that of the nominal performance of the multi-string GNC system, intended to return the crew to the earth in a survivable but unguided manner. Finally, various analyses will be discussed, including those completed to support validation efforts of the current CEV requirements, along with those on-going and planned with the intention to further refine the requirements and to support design development work in conjunction with the prime contractor. Some of these ongoing analyses will include work to size effectors (jets) and fuel budgets, to refine skip entry concepts, to characterize navigation performance and uncertainties, to provide for SM disposal offshore and to identify requirements to support target site selection.

Munday, S.↗
NASA NTRS2009

Orbiter Entry Aerodynamics Flight Testing: STS-119 and STS-125

NASA has performed entry flight testing related to boundary layer transition and turbulent heating environments during 2009. Two projects are involved in implementing the activities and acquiring flight data: 1) Orbiter BLT Flight Experiment during STS-119; and 2) Hypersonic Thermodynamic Infrared Measurements (HYTHIRM) during STS-119 and STS-125. Orbiter BLT FE has implemented tile surface thermocouples in order to provide in-situ data downstream of a fixed geometry tile protuberance. HYTHIRM has developed a framework of mission planning and aircraft based telescopic infrared measurements to provide quantitative surface temperature distributions.

Campbell, Charles H.↗
NASA NTRS2014

Mars Sample Return Using Commercial Capabilities: Propulsive Entry, Descent, and Landing of a Capsule Form Vehicle

This paper describes a critical portion of the work that has been done at NASA, Ames Research Center regarding the use of the commercially developed Dragon capsule as a delivery vehicle for the elements of a high priority Mars Sample Return mission. The objective of the investigation was to determine entry and landed mass capabilities that cover anticipated mission conditions. The "Red Dragon", Mars configuration, uses supersonic retro-propulsion, with no required parachute system, to perform Entry, Descent, and Landing (EDL) maneuvers. The propulsive system proposed for use is the same system that will perform an abort, if necessary, for a human rated version of the Dragon capsule. Standard trajectory analysis tools are applied to publically available information about Dragon and other legacy capsule forms in order to perform the investigation. Trajectory simulation parameters include entry velocity, flight path angle, lift to drag Ratio (L/D), landing site elevation, atmosphere density, and total entry mass, in addition engineering assumptions for the performance of the propulsion system are stated. Mass estimates for major elements of the overall proposed architecture are coupled to this EDL analysis to close the overall architecture. Three synodic launch opportunities, beginning with the 2022 opportunity, define the arrival conditions. Results state the relations between the analysis parameters as well as sensitivities to those parameters. The EDL performance envelope includes landing altitudes between 0 and -4 km referenced to the Mars Orbiter Laser Altimeter datum as well as minimum and maximum atmosphere density. Total entry masses between 7 and 10 mt are considered with architecture closure occurring between 9.0 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 and grow allowance for a Mars Ascent Vehicle (MAV), Earth Return Vehicle (ERV), and mission unique equipment. The useful payload supports an architecture that receives a sample from another surface asset and sends it directly back to Earth for recovery in a high Earth orbit. The work shows that emerging commercial capabilities as well as previously studied EDL methodologies can be used to efficiently support an important planetary science objective. The work also has applications for human exploration missions that will also use propulsive EDL techniques

EDL↗
NASA NTRSDate not supplied

Aerodynamic Performance of the Low-Earth Orbit Flight Test of an Inflatable Decelerator (LOFTID) Technology Demonstration Mission

The Low-Earth Orbit Flight Test of an Inflatable Decelerator (LOFTID) mission successfully demonstrated the deployment and entry performance of a 6 m diameter Hypersonic Inflatable Decelerator on November 10, 2022. This was the largest blunt body flown to date and demonstrated this inflatable aeroshell technology at scale and conditions relevant to Earth and Mars Entry, Descent, and Landing applications. LOFTID built upon the prior successes of the suborbital IRVE and IRVE-3 missions, leveraging and expanding on the flight experience from these and other planetary and Earth sample return missions. Both the re-entry vehicle and ejectable data module were successfully recovered from the Pacific Ocean, with LOFTID satisfying requirements for decelerator performance and stable flight from orbital entry conditions through parachute deployment and splashdown. This work discusses the pre-flight aerodynamics database and reconstructed flight performance of the LOFTID re-entry vehicle.

LOFTID↗
NASA NTRSDate not supplied

Aerodynamic Performance of the Low-Earth Orbit Flight Test of an Inflatable Decelerator (LOFTID) Technology Demonstration Mission

The Low-Earth Orbit Flight Test of an Inflatable Decelerator (LOFTID) mission successfully demonstrated the deployment and entry performance of a 6 m diameter Hypersonic Inflatable Decelerator on November 10, 2022. This was the largest blunt body flown to date and demonstrated this inflatable aeroshell technology at scale and conditions relevant to Earth and Mars Entry, Descent, and Landing applications. LOFTID built upon the prior successes of the suborbital IRVE and IRVE-3 missions, leveraging and expanding on the flight experience from these and other planetary and Earth sample return missions. Both the re-entry vehicle and ejectable data module were successfully recovered from the Pacific Ocean, with LOFTID satisfying requirements for decelerator performance and stable flight from orbital entry conditions through parachute deployment and splashdown. This work discusses the pre-flight aerodynamics database and reconstructed flight performance of the LOFTID re-entry vehicle.

LOFTID↗