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All sourcesNASA NTRSDOE OSTI
NASA NTRS2016

Thermophysics Issues Relevant to High-Speed Earth Entry of Large Asteroids

Physics of atmospheric entry of meteoroids was an active area of research at NASA ARC up to the early 1970s (e.g., the oft-cited work of Baldwin and Sheaffer). However, research in the area seems to have ended with the Apollo program, and any ties with an active international meteor physics community seem to have significantly diminished thereafter. In the decades following the 1970s, the focus of entry physics at NASA ARC has been on improvement of the math models of shock-layer physics (especially in chemical kinetics and radiation) and thermal response of ablative materials used for capsule heatshields. With the overarching objectives of understanding energy deposition into the atmosphere and fragmentation, could these modern analysis tools and processes be applied to the problem of atmospheric entry of meteoroids as well? In the presentation we will explore: (i) the physics of atmospheric entries of meteoroids using our current state-of-the-art tools and processes, (ii) how multiple bodies interact, and (iii) the influence of wall blowing on flow dynamics.

Atmospheric Entry↗
NASA NTRS2016

Entry at Venus

This is lecture to be given at the IPPW 2016, as part of the 2 day course on Short Course on Destination Venus: Science, Technology and Mission Architectures. The attached presentation material is intended to be introduction to entry aspects of Venus in-situ robotic missions. The presentation introduces the audience to the aerodynamic and aerothermodynamic aspects as well as the loads, both aero and thermal, generated during entry. The course touches upon the system design aspects such as TPS design and both high and low ballistic coefficient entry system concepts that allow the science payload to be protected from the extreme entry environment and yet meet the mission objectives.

Thermal protection↗
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 NTRS2019

Overview of Heatshield for Extreme Entry Environment Technology (HEEET)

The objective of the Heatshield for Extreme Entry Environment Technology (HEEET) projects is to mature a 3-D Woven Thermal Protection System (TPS) to Technical Readiness Level (TRL) 6 to support future NASA missions to destinations such as Venus and Saturn. Destinations that have extreme entry environments with heat fluxes > 3500 W/cm2 and pressures up to 5 atmospheres, entry environments that NASA has not flown since Pioneer-Venus and Galileo. The scope of the project is broad and can be split into roughly four areas, Manufacturing/Integration, Structural Testing and Analysis, Thermal Testing and Analysis and Documentation. Manufacturing/Integration covers from raw materials, piece part fabrication to final integration on a 1-meter base diameter 45-degree sphere cone Engineering Test Unit (ETU). A key aspect of the project was to transfer as much of the manufacturing technology to industry in preparation to support future mission infusion. The forming, infusion and machining approaches were transferred to Fiber Materials Inc. and FMI then fabricated the piece parts from which the ETU was manufactured. The base 3D-woven material consists of a dual layer weave with a high-density outer layer to manage recession in the system and a lower density, lower thermal conductivity inner layer to manage the heat load. At the start of the project it was understood that due to weaving limitations the heat shield was going to be manufactured from a series of tiles. And it was recognized that the development of a seam solution that met the structural and thermal requirements of the system was going to be the most challenging aspect of the project. It was also recognized that the seam design would drive the final integration approach and therefore the integration of the ETU was kept in-house within NASA. A final seam concept has been successfully developed and implemented on the ETU. The structural testing and analysis covers from characterization of the different layers of the infused material as functions of weave direction and temperature to sub-component level testing such as 4pt bend testing at sub-ambient and elevated temperature and culminates in testing of the ETU results from which will validate the structural models developed using the element and sub-component level tests. Given the seam has to perform both structurally and aerothermally during entry a novel 4pt bend test fixture was developed allowing articles to be tested while the front surface is heated with a laser. These tests are being utilized to establish the systems structural capability during entry.A broad range of aerothermal tests (arcjet tests) were performed to develop material response models for predicting the required TPS thickness to meet a missions needs and to evaluate failure modes and establish the capability of the system. The final aspect of the project is to develop a comprehensive Design and Data Book such that a future mission will have the information necessary to adopt the technology. This presentation will provide an overview for each of these areas and argue that HEEET has successfully achieved TRL 6.

HEEET↗
NASA NTRSDate not supplied

In Situ Small Spacecraft Missions Utilizing Heatshield for Extreme Entry Environments Technology

There is considerable interest in utilizing Small Spacecraft beyond low Earth orbit. In November of 2018, successful data relay operations of the MarCO CubeSats during the entry, descent, and landing (EDL) of the Mars InSight mission showed the viability of using CubeSats for interplanetary missions. Additional testament to the promise of Small Spacecraft class rideshare missions is the upcoming Artemis-1 flight test that will launch thirteen 6U CubeSats, as well as the establishment of NASA’s SIMPLEx program which will conduct stand-alone planetary science missions that launch with a primary payload. It is anticipated that continued innovations in Small Spacecraft capabilities combined with small EDL systems will expand the range of potential missions to allow for in situ investigations. Recently, NASA invested in the development of a new, efficient and capable ablative thermal protection system (TPS), utilizing 3-D Weaving. The new dual layer TPS, Heatshield for Extreme Entry Environment Technology (HEEET), is tailorable, scalable, robust, mass efficient and capable of supporting in situ missions across the solar system. Combining the HEEET entry system with innovative Small Spacecraft technology will substantially expand the range of Small Spacecraft mission applications by offering the capability for high speed entry or aerocapture at destinations with atmospheres. This paper will describe the HEEET aeroshell system and highlight various mission concepts including a dual technology demonstration mission that is under development and other concepts to deliver instruments for planetary science.

Small Spacecraft↗
NASA NTRSDate not supplied

Venus Cloud Layer Investigation: Aeroshells for Entry, Descent and Deployment

Entry, Descent and Deployment (EDD) of aerial platforms at Venus follows similar operational approach as landers. •Limited only by the availability of mass efficient and robust aeroshell (heatshield/TPS) technology. •Heatshield for Extreme Entry Environment Technology (HEEET) at TRL 6 is an enabler of Venus in-situ missions.Lower ballistic coefficient, deployable concept, ADEPT, offers additional options•Low deceleration entry profile•Release of one or more payloads (balloons) from open back of the entry vehicle2

Venus↗
NASA NTRSDate not supplied

Impact of NASA’s Entry Systems Modeling Project on Planetary Mission Design

Planetary missions continue to grow larger and more complex. Furthermore, the current focus on human exploration of the Moon and Mars, as well as Mars Sample Return(MSR), place increas-ingly stringent requirements on the reliability of the entry, descent, and landing (EDL) system that ensures the safe delivery of payload or crew to their destination. Planetary EDL is an area in which mission designers are critically reliant on modeling and simulation to demonstrate the reliability of the system, as there are no ground facilities that are able to fully test these systems in a flight-relevant environment. NASA’s state-of-the-art modeling and simulation capability must continually evolve to meet the needs of the next generation of planetary EDL. To accomplish this aim, NASA’s Entry Systems Modeling (ESM) Project was formed in 2013and is funded bythe Space Technology Mission Directorate(STMD) and Science Mission Directorate (SMD). ESM is the Agency’s only cross-cutting effort for advancing entry systems modeling and simulation capabilities across a range of technical disciplines and Solar System destinations. ESM is a portfolio project covering a variety of mid-TRL research efforts within four core EDL-related areas of investment: (1) Thermal protection material modeling, (2) Shock layer kinetics and radiation, (3) Aerosciences, and (4) Guidance, navigation, and con-trol. The material modeling group creates detailed material response modelsof thermal protection systems (TPS)from the micro to macro scale, and at the fun-damental and engineering levels. Shock layer kinetics and radiation focuses on radiative heating of space-craft, quantum chemistry and benchmark experiments for validation.Aerosciences is a broad research area that impacts many aspects of entry systems, including parachutes, aerodynamics, and turbulent heating augmentation due to TPS roughness.The guidance, navigation,and control effort under ESM is expanding the capabilities of NASA’s main flight mechanics tool, POST2, for use on high-performance computing architectures and to generalize interoperability with external applications for more detailed end-to-end simula-tion.In addition, several focusedresearch topics have been approvedto augment ESM’s core portfolio. These include efforts for deep post-flight analysis of Mars 2020/MEDLI2 flight data; development ofTPS failure models; improvement of hypersonic wakeflow models; and a recently concluded effort to provide material response models for NuSil-coated PICA heat-shield material. This presentation will discuss each of these investment areas and demonstrate via real mission examples how advances to the state-of-the-art enabled by ESM are directly impacting the missions of today and tomorrow, including InSight, Mars 2020, Mars Sample Return, Orion, and Dragonfly.

Entry Systems Modeling↗
NASA NTRSDate not supplied

Design and Technology Maturation of the Stratospheric Projectile Experiment of Entry Dynamics

The supersonic and transonic dynamic stability of blunt-body reentry vehicles currently poses large risks in all of NASA’s ongoing entry missions (MSR SRL, MSR EES, and Dragonfly). These projects have allocated millions of dollars to testing and modeling efforts to buy down risk by using the current state-of-the-art (SoA) facilities at NASA’s disposal. While these facilities have heritage in supplying dynamics data to reentry missions, their availability is severely limited – particularly with the high number of concur-rent projects requesting simultaneous testing– and are costly when considering the science density per dollar. None of the current SoA facility methodologies allow the test model to have the dynamics fully develop through a flight relevant free-stream profile and as such require extrapolations with resultant high uncertainties in order to relate the test dynamics to flight expectations. SPEED is a NASA Ames Center Innovation Fund (CIF) project that is developing a highly tailorable and cost-effective test methodology to better assess the dynamic stability of blunt-body reentry vehicles via a stratospheric balloon flight. This is accomplished by dropping a suite of instrumented capsules from a stratospheric balloon to gain a statistically relevant dataset of scaled reentry vehicles in mission relevant free-flight conditions. This presentation will walk through how the test methodology is being implemented specifically for the Mars Sample Return (MSR) Earth Entry System (EES) geometry in an awarded Flight Opportunities Program (FOP) test flight in early CY24. SPEED Application to MSR: SPEED consists of three main mechanical systems: the Drop Platform, the Projectile, and the test Capsule. SPEED is being developed as a set of guidelines and recommendations for how to test with the proposed Concept of Operations (Conops) since the specific design parameters will vary depending on the specific project’s reference trajectory and entry vehicle design. As such, this presentation will walk through the development time-line as shown in Fig. 2. This is meant to serve as a blueprint for further missions as desired. Mechanical and Avionics Design. The SPEED test platform designed for the MSR-EES capsule geometry with nominal entry parameters has the ability to carry 10 Capsules to altitude instrumented with: 1. 3-Axis Accelerometer 2. IMU 3. Gyroscope 4. Magnetometer 5. Pressure Transducer cruciform 6. Uplook and Horizon Cameras To package the avionics/instrumentation suite, the capsule is approximately 1’ in diameter with the Outer Mold Line (OML) centroid-scaled from the full EES design. The internal volume is gutted and custom-shaped to fit the desired instrumentation suite, as well as to allow for the positioning of ballast mass such that the Center of Gravity is analogous to the flight vehicle. All structural components in the Capsule and Projectile are 3D printed, which significantly reduces the cost of each flight unit to around $1500 including all instrumentation, avionics, and structural components. Flight Conops. The test Capsule is accelerated to the desired altitude and Mach number while stowed in the Projectile, a missile-like vehicle consisting of steel ballast in the nose, a low-drag OML, and an Ejection Mechanism to reliably release the Capsule into the free-flow supersonic conditions. For the MSR-EES design, the capsule employs ~3kg of ballast mass at the nose to accelerate the 1.25kg test Capsule to ~Mach 1.7 at 23km altitude. This requires an initial release altitude of 40km, the quoted limit of a 80kg payload by the FOP-contracted balloon provider. Once the Ejection Mechanism avionics detect the proper conditions, the spring-loaded Ejection Mechanism will release and – guided by the sabot – expose the test Capsule to the desired test conditions for ~5 seconds of free-flight in the supersonic/transonic regimes. Dynamics in the subsonic regime will also be captured with the instrumentation suite with post-flight recovery operations aimed at recovering the high-G-load capable SD cards after the planned hard impact landings. Testing and Development: In the few months the SPEED project has worked the development of MSR-EES flight test, the team has performed lab and drone based testing which this presentation will overview. After the first design phase, the team fabricated Engineering Demonstration Units (EDUs) of all subsystems to perform validation testing shown in Fig. 5. After validation was completed on the subsystem level, a drone-drop test was performed at the recreational flight ceiling of 400ft altitude to assess the SPEED systems in a flight environment. Parameters such as in-flight stability, hard impact landing performance, and avionics performance were quantified and qualified. The FY23 CIF will culminate in a helicopter drop test aboard an Air National Guard Blackhawk. This will prepare the team for the CY24 FOP stratospheric balloon flight that should provide the final verification to begin offering the test platform for mission support. Focus of Presentation: This presentation will outline the technology maturation path of the SPEED implementation to the MSR-EES capsule baseline as well as the details regarding the mechanical system, avionics and instrumentation, and flight operations. Note that a complementary presentation is being submitted for a methodology overview of the SPEED test platform, introducing the testing technique and benefits as well as the full application space of the technology.

pitch damping coefficient↗
NASA NTRSDate not supplied

Development of a Low-Cost, Highly Flexible Re-Entry Platform

With a growing interest in atmospheric entry, descent, and landing research as well as the necessity to empirically test increasingly complicated entry systems, a growing need for a quick and cost-effective method of experimentation is needed. The Hypersonic Configurable Unit Ballistic Experiment is a test and evaluation platform allowing for in-situ research and rapid testing of novel technologies in a re-entry environment. HyCUBE takes the form of a small satellite-sized re-entry vehicle and, with the additional space on board the vehicle, can act as a platform for testing novel guidance and control methodologies.

Aero-Thermodynamics↗
NASA NTRSDate not supplied

Assessing Huygens Probe Entry, Descent, and Landing at Titan Simulation using Dragonfly Atmosphere Model

Dragonfly is a New Frontiers Program mission that will deliver a rotorcraft to Saturn's moon, Titan. This mission follows Huygens as the previous mission that successfully landed a vehicle on Titan. A flight mechanics simulation of Dragonfly's Entry, Descent, and Landing sequence has been developed using the Program to Optimize Simulated Trajectories II. The simulation incorporates several subsystem models, including aerodynamics, gravity, and mass properties, to fully capture the multi-body six degree of freedom dynamics. Among all the subsystem models that inform the Entry, Descent, and Landing dynamics, the atmosphere model of Titan is a critical component. The atmosphere model characterizes the density, temperature, pressure, and winds that the entry vehicle experiences during the descent. This impacts several aspects of the descent such as the peak heating, aerodynamics, parachute release conditions, the dynamics of the vehicle and parachutes, and the landing ellipse. In the course of developing Dragonfly, an updated model of the Titan atmosphere has been created corresponding to Dragonfly's arrival in the mid-2030s, approximately one Titan year after the Huygens mission successfully landed a probe on Titan. This work leverages previous work done to investigate Huygens EDL sequence to assess the atmosphere model developed for Dragonfly. This is done by utilizing the updated Titan atmosphere model, the Dragonfly atmosphere model, within the Huygens POST2-based flight simulation with the goal of characterizing the differences between the atmospheric models and assessing how the Dragonfly atmosphere model impacts Huygens entry dynamics.

Huygens↗
NASA NTRSDate not supplied

Dragonfly Mission Entry and Descent Modeling and Simulation Overview

Dragonfly is a New Frontiers Program mission, led by The Johns Hopkins Applied Physics Laboratory, that will deliver a rotorcraft lander to Saturn’s moon, Titan. The focus of this work is to analyze the trajectory of the entry vehicle from cruise stage separation until lander separation. This analysis is done by the NASA Langley Research Center Entry, Descent, and Landing team using the Program to Optimize Trajectories II. This paper provides an overview of the current design and the robustness of the overall entry sequence using a Monte Carlo uncertainty analysis. The work presented in this study includes the updated design, models, and analysis completed since the Dragonfly Entry, Descent, and Landing Mission Preliminary Design Review.

Simulation↗
NASA NTRSDate not supplied

Assessing Huygens Probe Entry, Descent, and Landing at Titan Simulation using Dragonfly Atmosphere Model

Dragonfly is a New Frontiers Program mission that will deliver a rotorcraft to Saturn's moon, Titan. This mission follows Huygens as the previous mission that successfully landed a vehicle on Titan. A flight mechanics simulation of Dragonfly's Entry, Descent, and Landing sequence has been developed using the Program to Optimize Simulated Trajectories II. The simulation incorporates several subsystem models, including aerodynamics, gravity, and mass properties, to fully capture the multi-body six degree of freedom dynamics. Among all the subsystem models that inform the Entry, Descent, and Landing dynamics, the atmosphere model of Titan is a critical component. The atmosphere model characterizes the density, temperature, pressure, and winds that the entry vehicle experiences during the descent. This impacts several aspects of the descent such as the peak heating, aerodynamics, parachute release conditions, the dynamics of the vehicle and parachutes, and the landing ellipse. In the course of developing Dragonfly, an updated model of the Titan atmosphere has been created corresponding to Dragonfly's arrival in the mid-2030s, approximately one Titan year after the Huygens mission successfully landed a probe on Titan. This work leverages previous work done to investigate Huygens EDL sequence to assess the atmosphere model developed for Dragonfly. This is done by utilizing the updated Titan atmosphere model, the Dragonfly atmosphere model, within the Huygens POST2-based flight simulation with the goal of characterizing the differences between the atmospheric models and assessing how the Dragonfly atmosphere model impacts Huygens entry dynamics.

Entry Descent Landing↗
NASA NTRSDate not supplied

Dragonfly Mission Entry and Descent Modeling and Simulation Overview

Dragonfly is a New Frontiers Program mission, led by The Johns Hopkins Applied Physics Laboratory, that will deliver a rotorcraft lander to Saturn’s moon, Titan. The focus of this work is to analyze the trajectory of the entry vehicle from cruise stage separation until lander separation. This analysis is done by the NASA Langley Research Center Entry, Descent, and Landing team using the Program to Optimize Trajectories II. This paper provides an overview of the current design and the robustness of the overall entry sequence using a Monte Carlo uncertainty analysis. The work presented in this study includes the updated design, models, and analysis completed since the Dragonfly Entry, Descent, and Landing Mission Preliminary Design Review.

Simulation↗
NASA NTRSDate not supplied

AI-Enhanced Computational Tools for Entry Systems Modeling

To advance the understanding of complex atmospheric entry phenomena, NASA’s Entry Systems Modeling (ESM) team [1] has developed high-fidelity computational tools addressing multiscale challenges, from material microstructures to full-scale heatshield response. This abstract highlights a subset of ESM tools, focusing on AI integration to enhance workflows and predictive modeling. - PuMA [2] computes effective material properties from high-resolution micro-CT scans, supporting TPS analysis for NASA missions. - TomoSAM [3] automates 3D tomography dataset segmentation for PuMA using the Segment Anything Model, reducing manual effort and improving accuracy. - PATO [4] models porous reactive materials under extreme conditions, with advancements such as unified solvers, mechanical erosion, and TPS coatings for NASA missions. - arcjetCV [5] employs deep learning to analyze arc jet test footage, measuring recession rates, shape changes, and shock standoff distances, bridging simulations, and experiments to reveal TPS ablation behavior. - ARCHeS [6] simulates arc heater plasma flows, modeling turbulence, radiation, and electromagnetic interactions to optimize arc heater performance, validate TPS under extreme conditions, and serve as a foundation for developing digital twins of arc heater facilities. - SPARTA [7] simulates rarefied hypersonic flows and gas-surface interactions for planetary entry missions, leveraging GPU architectures for scalable and efficient aerothermal and ablation analyses. AI-driven solutions, such as deep learning segmentation, have streamlined workflows in ESM tools and still hold significant potential to further accelerate processes and enhance automation in entry systems modeling. [1] Haskins, J.B. (2023), [2] Ferguson, J.C. (2018), [3] Meurisse, J.B.E. (2018), [4] Semeraro, F. (2023), [5] Quintart, A. (2024) [6] Meurisse, J.B.E. (2022), [7] Plimpton, S.J. (2019)

Predictive Modeling↗
DOE OSTI2025

Acoustic Observations of the OSIRIS-REx Sample Return Capsule Re-Entry from Wendover Airport

The Origins, Spectral Interpretation, Resource Identification, and Security‐Regolith Explorer sample return capsule (SRC) re‐entered the Earth’s atmosphere at hypersonic speeds from interplanetary space on 24 September 2023. The current work reports on 18 ground‐based acoustic sensors deployed at Wendover Airport, the same location that the Genesis and Stardust SRC re‐entries were recorded. Four different sensors (Chaparral Physics, Gem, Wilson Engineering Research and Development [WERD], and RedVox) were deployed in close proximity to compare their performance. All the sensors captured an N‐wave signal associated with the SRC re‐entry shock wave followed by a broadband coda. The Chaparral Physics array served as the high‐fidelity reference measurement. The N‐wave signal had a peak‐to‐peak amplitude of 4.07 Pa with a fundamental frequency of 4.98 Hz from 167.5° measured clockwise from north, nearly perpendicular to the SRC trajectory. In addition, high coherence in the coda was shown to be associated reflections from the surrounding mountains. In general, the more economical sensors (Gem, WERD, and RedVox) produced results that were consistent with these observations and sensor specifications. Beamforming with these single sensors arranged as an array showed agreement with the high‐fidelity array to within a couple of degrees. Furthermore, the current high‐fidelity results were compared with the measurements during the Genesis and Stardust SRC re‐entries. All three entries produced a broadband fundamental peak at a frequency that was inversely related to the SRC diameter as well as evidence of reflections from the surrounding topography.

KC, Real J. [Oklahoma State University, Stillwater↗
NASA NTRS1972

Apollo experience report: Mission planning for Apollo entry

The problems encountered and the experience gained in the entry mission plans, flight software, trajectory-monitoring procedures, and backup trajectory-control techniques of the Apollo Program should provide a foundation upon which future spacecraft programs can be developed. Descriptions of these entry activities are presented. Also, to provide additional background information needed for discussion of the Apollo entry experience, descriptions of the entry targeting for the Apollo 11 mission and the postflight analysis of the Apollo 10 mission are presented.

Graves, C. A.↗
NASA NTRS1972

Separation and communications geometry analysis for a Jupiter entry probe from a Pioneer spacecraft

The separation and communications geometry analysis is presented which shows that it is possible to target a Jupiter entry probe from a Pioneer spin-stabilized vehicle to desired entry conditions and maintain a continuous line-of-sight communications link while the probe descends to a pressure of 10 bars within the atmosphere. Such a targeting is possible with either the deflected probe mode or with the deflected bus mode. For the nominal mission studied, an 800 day trip in 1978 with the Pioneer type spacecraft flying by a 1.3 R(J), it was possible to achieve entries near zero angle of attack at about -25 deg relative entry flight path angle. Because of the constraint on the deflection maneuver for the deflected probe mode, the deflection and phasing velocity requirements are much larger for this mode. The geometry parameters, look angles, ranges and range rates, provided by this analysis permit the design of a functional communications system for the bus and probe. Avoiding a high trapped radiation belt hazard at Jupiter by flying the bus by at 6 R(J), however, requires using a despun antenna on the flyby bus.

Tindle, E. L.↗
NASA NTRS1973

Structure of the ionospheric disturbances about planetary entry probes

Local ionospheric disturbances which would be created by a planetary entry probe are investigated. Competing theories of spacecraft-ambient plasma interactions are used to estimate computationally the perturbations of the plasma, particularly the structure of the near wake behind planetary entry vehicles. The results have bearing on the location and operation of plasma diagnostic instrumentation aboard planetary entry vehicles. Recent estimates of Mars ionospheric properties plus vehicle dimensions and speeds similar to those of the Viking Mars Lander are used to define the parameters essential to the theory. Smaller entry bodies are also considered. Comparisons are made of the results based on the different theories for a given assumed planetary atmosphere, and also with the perturbations a similar vehicle would generate in the earth's ionosphere.

Weil, H.↗