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NASA NTRS2023

Dragonfly: A Case Study in Dynamic Stability Characterization

Next decade, the Dragonfly probe will enter Titan’s atmosphere on a mission to explore and characterize its unique and complex organic chemistry. In contrast to Mars entry missions, where the challenge is to slow down in its tenuous atmosphere before impacting the surface, Titan’s thick atmosphere provides generous drag. However, Titan has its own unique EDL challenges. Among these challenges is the long time (on the order of 90 minutes!) the vehicle will need to descend unguided through flow regimes where blunt-body entry probes are known to be dynamically unstable. This elongated timeline, both prior to drogue parachute deployment, and during the ensuing descent, has the potential to allow adverse dynamic instability properties to produce large amplitude oscillations in angle-of-attack. This, in-turn may lead to non-compliance in vehicle attitude requirements for EDL events such as parachute deployment, and lander release, or in the most extreme case, tumbling of the aeroshell and near-assured loss of mission.

EDL↗
NASA NTRS2023

Dragonfly: A Case Study in Dynamic Stability Characterization

Next decade, the Dragonfly probe will enter Titan’s atmosphere on a mission to explore and characterize its unique and complex organic chemistry. In contrast to Mars entry missions, where the challenge is to slow down in its tenuous atmosphere before impacting the surface, Titan’s thick atmosphere provides generous drag. However, Titan has its own unique EDL challenges. Among these challenges is the long time (on the order of 90 minutes!) the vehicle will need to descend unguided through flow regimes where blunt-body entry probes are known to be dynamically unstable. This elongated timeline, both prior to drogue parachute deployment, and during the ensuing descent, has the potential to allow adverse dynamic instability properties to produce large amplitude oscillations in angle-of-attack. This, in-turn may lead to non-compliance in vehicle attitude requirements for EDL events such as parachute deployment, and lander release, or in the most extreme case, tumbling of the aeroshell and near-assured loss of mission.

EDL↗
NASA NTRS1990

Entry vehicle performance analysis and atmospheric guidance algorithm for precision landing on Mars

Future missions to Mars may require pin-point landing precision, possibly on the order of tens of meters. The ability to reach a target while meeting a dynamic pressure constraint to ensure safe parachute deployment is complicated at Mars by low atmospheric density, high atmospheric uncertainty, and the desire to employ only bank angle control. The vehicle aerodynamic performance requirements and guidance necessary for 0.5 to 1.5 lift drag ratio vehicle to maximize the achievable footprint while meeting the constraints are examined. A parametric study of the various factors related to entry vehicle performance in the Mars environment is undertaken to develop general vehicle aerodynamic design requirements. The combination of low lift drag ratio and low atmospheric density at Mars result in a large phugoid motion involving the dynamic pressure which complicates trajectory control. Vehicle ballistic coefficient is demonstrated to be the predominant characteristic affecting final dynamic pressure. Additionally, a speed brake is shown to be ineffective at reducing the final dynamic pressure. An adaptive precision entry atmospheric guidance scheme is presented. The guidance uses a numeric predictor-corrector algorithm to control downrange, an azimuth controller to govern crossrange, and analytic control law to reduce the final dynamic pressure. Guidance performance is tested against a variety of dispersions, and the results from selected tests are presented. Precision entry using bank angle control only is demonstrated to be feasible at Mars.

Dieriam, Todd A.↗
NASA NTRS2022

A Rideshare Tensegrity Rover Concept to Explore Titan’s Lands and Oceans

The Tension Adjustable Network for Deploying Entry Membrane (TANDEM) concept is a phase II NASA Innovative Advanced Concepts (NIAC) project. One of the goals for this project was to extend the application of tensegrity-based rovers to a rideshare concept for a Titan mission. This report provides preliminary details of the TANDEM concept for Titan. System analyses of the entry vehicle were performed and compared to data from the Huygens mission to Titan. The entry aeroshell was based on Huygens and Stardust aeroshells. Modeling, simulation and design of the tensegrity deployment were performed using a nonlinear optimization form-finding algorithm. Lastly, a neuroevolution-based machine learning control strategy was applied, which produced efficient rolling locomotion gaits.

tensegrity↗
NASA NTRS2009

Supersonic Testing of 0.8 m Disk Gap Band Parachutes in the Wake of a 70 Deg Sphere Cone Entry Vehicle

Supersonic wind tunnel testing of Viking-type 0.8 m Disk-Gap-Band (DGB) parachutes was conducted in the NASA Glenn Research Center 10'x10' wind-tunnel. The tests were conducted in support of the Mars Science Laboratory Parachute Decelerator System development and qualification program. The aerodynamic coupling of the entry-vehicle wake to parachute flow-field is under investigation to determine the cause and functional dependence of a supersonic canopy breathing phenomenon referred to as area oscillations, characteristic of DGB's above Mach 1.5 operation. Four percent of full-scale parachutes (0.8 m) were constructed similar to the flight-article in material and construction techniques. The parachutes were attached to a 70-deg sphere-cone entry-vehicle to simulate the Mars flight configuration. The parachutes were tested in the wind-tunnel from Mach 2 to 2.5 in a Reynolds number range of 2x105 to 1x106, representative of a Mars deployment. Three different test configurations were investigated. In the first two configurations, the parachutes were constrained horizontally through the vent region to measure canopy breathing and wake interaction for fixed trim angles of 0 and 10 degrees from the free-stream. In the third configuration the parachute was unconstrained, permitted to trim and cone, similar to free-flight (but capsule motion is constrained), varying its alignment relative to the entry-vehicle wake. Non-intrusive test diagnostics were chosen to quantify parachute performance and provide insight into the flow field structure. An in-line loadcell provided measurement of unsteady and mean drag. Shadowgraph of the upstream parachute flow field was used to capture bow-shock motion and wake coupling. Particle image velocimetry provided first and second order flow field statistics over a planar region of the flow field, just upstream of the parachute. A photogrammetric technique was used to quantify fabric motion using multiple high speed video cameras to record the location in time and space of reflective targets placed on the canopy interior. The experimental findings including an updated drag model and the physical basis of the area oscillation phenomenon will be discussed.

EDL↗
NASA NTRS2013

Orion Exploration Flight Test-1 Contingency Drogue Deploy Velocity Trigger

As a backup to the GPS-aided Kalman filter and the Barometric altimeter, an "adjusted" velocity trigger is used during entry to trigger the chain of events that leads to drogue chute deploy for the Orion Multi-Purpose Crew Vehicle (MPCV) Exploration Flight Test-1 (EFT-1). Even though this scenario is multiple failures deep, the Orion Guidance, Navigation, and Control (GN&C) software makes use of a clever technique that was taken from the Mars Science Laboratory (MSL) program, which recently successfully landing the Curiosity rover on Mars. MSL used this technique to jettison the heat shield at the proper time during descent. Originally, Orion use the un-adjusted navigated velocity, but the removal of the Star Tracker to save costs for EFT-1, increased attitude errors which increased inertial propagation errors to the point where the un-adjusted velocity caused altitude dispersions at drogue deploy to be too large. Thus, to reduce dispersions, the velocity vector is projected onto a "reference" vector that represents the nominal "truth" vector at the desired point in the trajectory. Because the navigation errors are largely perpendicular to the truth vector, this projection significantly reduces dispersions in the velocity magnitude. This paper will detail the evolution of this trigger method for the Orion project and cover the various methods tested to determine the reference "truth" vector; and at what point in the trajectory it should be computed.

Gay, Robert S.↗
NASA NTRS2007

Overview of the Mars Science Laboratory Parachute Decelerator Subsystem

In 2010 the Mars Science Laboratory (MSL) mission will deliver NASA's largest and most capable rover to the surface of Mars. MSL will explore previously unattainable landing sites due to the implementation of a high precision Entry, Descent, and Landing (EDL) system. The parachute decelerator subsystem (PDS) is an integral prat of the EDL system, providing a mass and volume efficient some of aerodynamic drag to decelerate the entry vehicle from Mach 2 to subsonic speeds prior to final propulsive descent to the sutface. The PDS for MSL is a mortar deployed 19.7m Viking type Disk-Gap-Band (DGB) parachute; chosen to meet the EDL timeline requirements and to utilize the heritage parachute systems from Viking, Mars Pathfinder, Mars Exploration Rover, and Phoenix NASA Mars Lander Programs. The preliminary design of the parachute soft goods including materials selection, stress analysis, fabrication approach, and development testing will be discussed. The preliminary design of mortar deployment system including mortar system sizing and performance predictions, gas generator design, and development mortar testing will also be presented.

Entry, Descent, and Landing (EDL)↗
NASA NTRS2008

Supersonic Disk Gap Band Parachute Performance in the Wake of a Viking-Type Aeroshell from Mach 2 to 2.5

Supersonic wind tunnel testing of 0.813 m diameter Disk-Gap-Band parachutes is being conducted in the NASA Glenn Research Center (GRC) 10' x 10' wind-tunnel. The tests are conducted in support of the Mars Science Laboratory Parachute Decelerator System development and qualification. Four percent of full-scale parachutes were constructed similarly to the flight-article in material and construction techniques. The parachutes are attached to a 4% scale MSL entry-vehicle to simulate the free-flight configuration. The parachutes are tested from Mach 2 to 2.5 over a Reynolds number (Re) range of 1 to 3 x 10(exp 6), representative of the MSL deployment envelope. Constrained and unconstrained test configurations are investigated to quantify the effects of parachute trim, suspension line interaction, and alignment with the capsule wake. The parachute is constrained horizontally through the vent region, to measure canopy breathing and wake interaction for fixed trim angles of 0 and 10 degrees from the velocity vector. In the unconstrained configuration the parachute is permitted to trim and cone, similar to the free-flight varying its alignment relative to the entry-vehicle wake. Test diagnostics were chosen to quantify parachute performance and to provide insight into the flow field structure. An in-line load cell provided measurement of unsteady and mean drag as a function of Mach and Re. High-speed shadowgraph video of the upstream parachute flow field was used to capture bow-shock motion and stand of distance. Particle image velocimetry of the upstream parachute flow field provides spatially and temporally resolved measurement velocity and turbulent statistics. Multiple high speed video views of targets placed in the interior of the canopy enable photo-grammetric measurement of the fabric motion in time and space from reflective. High speed video is also used to document the supersonic inflation and measure trim angle, projected area, and frequency of area oscillations.

EDL↗
NASA NTRS2011

Flexible Ablators: Applications and Arcjet Testing

Flexible ablators were conceived in 2009 to meet the technology pull for large, human Mars Exploration Class, 23 m diameter hypersonic inflatable aerodynamic decelerators. As described elsewhere, they have been recently undergoing initial technical readiness (TRL) advancement by NASA. The performance limits of flexible ablators in terms of maximum heat rates, pressure and shear remain to be defined. Further, it is hoped that this emerging technology will vastly expand the capability of future NASA missions involving atmospheric entry systems. This paper considers four topics of relevance to flexible ablators: (1) Their potential applications to near/far term human and robotic missions (2) Brief consideration of the balance between heat shield diameter, flexible ablator performance limits, entry vehicle controllability and aft-body shear layer impingement of interest to designers of very large entry vehicles, (3) The approach for developing bonding processes of flexible ablators for use on rigid entry bodies and (4) Design of large arcjet test articles that will enable the testing of flexible ablators in flight-like, combined environments (heat flux, pressure, shear and structural tensile loading). Based on a review of thermal protection system performance requirements for future entry vehicles, it is concluded that flexible ablators have broad applications to conventional, rigid entry body systems and are enabling to large deployable (both inflatable and mechanical) heat shields. Because of the game-changing nature of flexible ablators, it appears that NASA's Office of the Chief Technologist (OCT) will fund a focused, 3-year TRL advancement of the new materials capable of performance in heat fluxes in the range of 200-600 W/sq. cm. This support will enable the manufacture and use of the large-scale arcjet test designs that will be a key element of this OCT funded activity.

Thermal Protection↗
NASA NTRS2014

Orion Multi-Purpose Crew Vehicle (MPCV) Capsule Parachute Assembly System (CPAS) Wake Deficit Wind Tunnel Testing

During descent after re-entry into the Earth's atmosphere, the Orion CM deploys its drogue parachutes at approximately Mach 0.7. Accurately predicting the dynamic pressure experienced by the drogue parachutes at deployment is critical to properly designing the parachutes. This NASA Engineering and Safety Center assessment was designed to provide a complete set of flowfield measurements on and around an idealized Orion Crew Module shape with the most appropriate wind tunnel simulation of the Orion flight conditions prior to parachute deployment. This document contains the details of testing and the outcome of the assessment.

Ross, James C.↗
NASA NTRSDate not supplied

Interpretation of Vehicle Tumbling Predictions from 6-DOF Entry and Descent Simulation

Blunt body entry vehicles are subject to dynamic instability during terminal descent. This often manifests as limit cycle oscillations in total angle of attack, but can diverge into tumbling behavior under certain conditions. For the Mars Sample Return Earth Entry Vehicle (MSR EEV), there is a constraint on the orientation of the sample tubes so backward impact is impermissible. Past missions have chosen to deploy parachutes to preclude tumbling, but active events after release of MSR EEV have been ground-ruled out with the intent to maximize system reliability. Prevention of tumbling during subsonic descent is a design driver for MSR EEV. During preliminary design of the MSR EEV, six degree-of-freedom numerical simulations indicated an unacceptably high probability of tumbling for a 60degree sphere-cone forebody geometry, which necessitated a design change. Decreasing the forebody angle was expected to improve dynamic stability, but would also adversely impact mass, aeroheating, and manufacturing risk. Hence there was strong motivation to understand the physical drivers for the onset of tumbling, and to determine: (a) whether the causes of tumbling are representative of physically realizable vehicle entry configurations and (b) what changes can be made to existing design and analysis practices to ensure a stable vehicle.

Entry Vehicle↗
NASA NTRS2017

Lightweight Multifunctional Planetary Probe for Extreme Environment Exploration and Locomotion

The demand to explore new worlds requires the development of advanced technologies that enable landed science on uncertain terrains or in hard to reach locations. As a result, contemporary Entry, Descent, Landing, (EDL) and additional locomotion (EDLL) profiles are becoming increasingly more complex, with the introduction of lifting/guided entries, hazard avoidance on descent, and a plethora of landing techniques including airbags and the skycrane maneuver. The inclusion of each of these subsystems into a mission profile is associated with a substantial mass penalty. This report explores the new all-in-one entry vehicle concept, TANDEM, a new combined EDLL concept, and compares it to the current state of the art EDL systems. The explored system is lightweight and collapsible and provides the capacity for lifting/guided entry, guided descent, hazard avoidance, omnidirectional impact protection and surface locomotion without the aid of any additional subsystems. This Phase I study explored: 1. The capabilities and feasibility of the TANDEM concept as an EDLL vehicle. 2. Extensive impact analysis to ensure mission success in unfavorable landing conditions, and safe landing in Tessera regions. 3. Development of a detailed design for a conceptual mission to Venus. As a result of our work it was shown that: 1. TANDEM provides additional benefits over the Adaptive, Deployable Entry Placement Technology (ADEPT) including guided descent and surface locomotion, while reducing the mass by 38% compared to the ADEPT-VITaL mission. 2. Demonstrated that the design of tensegrity structures, and TANDEM specifically, grows linearly with an increase in velocity, which was previously unknown. 3. Investigation of surface impact revealed a promising results that suggest a properly configured TANDEM vehicle can safely land and preform science in the Tessera regions, which was previously labeled by the Decadal Survey as, largely inaccessible despite its high scientific interest. This work has already resulted in a NASA TM and will be submitted to the Journal of Spacecraft and Rockets.

landing techniques↗
NASA NTRSDate not supplied

Evaluation of the LOFTID Flight Thermocouple Measurements

The Flexible Thermal Protection System (FTPS) on NASA’s Low-Earth Orbit Flight Test of an Inflatable Decelerator (LOFTID) vehicle was instrumented with thermocouples (TC) to measure the in-depth thermal response during entry into Earth’s atmosphere. Accurate flight temperature measurements are critical for verifying vehicle performance during the flight test and reducing uncertainties in the thermal models. The deployable nature of inflatable decelerator technology presents challenges, namely the TCs need to be compactable, cannot damage the FTPS nor the inflatable structure, and need to be able to withstand high temperatures and large thermal gradients over significant running lengths. As previously reported, an extensive ground-test campaign was conducted prior to flight to inform the selection of the LOFTID FTPS TCs which minimized all known potential error sources. Overall, the resulting flight FTPS TC measurements were exceptionally clean throughout the atmospheric entry heat pulse. However, a few of the FTPS TCs still exhibited electrical shorting or other anomalous behavior. This paper presents the flight measurements made by the LOFTID FTPS TCs, identifies anomalous behavior, and discusses potential explanations for the causes of the anomalous behavior. Lastly, recommendations to further mitigate TC measurement error on future Hypersonic Inflatable Aerodynamic Decelerator (HIAD) missions are provided.

Ruth A Miller↗
NASA NTRS2020

NASA’s Space Launch System: Launch Capability for Lunar Exploration and Transformative Science

Excitement is building for the first launch of NASA’s Space Launch System (SLS), a unique exploration asset for the agency’s Artemis lunar program as well as for a new generation of science missions. SLS is designed for an array of missions beyond Earth’s orbit. The flexible system, which can be configured for Orion, cargo or Orion with co-manifested payload missions, offers high escape velocities to send more mass to deep space destinations. When configured with an 8.4 m-diameter fairing, SLS offers unmatched payload volume for human exploration and science missions. The initial Block 1 variant will insert at least 26 metric tons (t) to trans-lunar injection (TLI) and the more powerful Block 1B vehicle will launch 34-37 t to TLI using a new-development upper stage. Much of the initial SLS Block 1 vehicle is complete, including the upper stage and payload section, the core stage, engines and the solid rocket boosters. The first mission, Artemis I, launching from modernized and upgraded facilities at Kennedy Space Center (KSC), will be an uncrewed test flight of SLS, Orion and ground processing, with a primary objective of testing Orion’s heat shield at lunar re-entry velocity. Artemis I will have accommodations for 13 6U CubeSat payloads. These CubeSat missions will be deployed along the upper stage disposal trajectory after Orion separates from the vehicle. A rare opportunity for CubeSats to be deployed beyond low Earth orbit (LEO), Artemis I CubeSat missions range from searching for hydrogen and other volatiles on the lunar South Pole to studying the acceleration mechanisms of solar and interplanetary particles from a heliocentric trajectory. With manufacturing of the initial vehicle complete, fabrication and procurement is progressing for the second flight of SLS and Orion, Artemis II. Also an SLS Block 1 and Orion flight launching from KSC, Artemis II will mark the return of American astronauts to deep space with a lunar flyby-free return trajectory mission. With the Artemis III flight, NASA has the goal to land the first woman and the next man on the Moon. Infrastructure beyond SLS will be required for this effort, including elements of the lunar Gateway as well as lunar rovers, landers and additional commercially supplied launch services. SLS, as the only vehicle with the capability to lift 26 t of mass to TLI in its initial Block 1 variant, will remain a key component of this new-era exploration program. Future variants – Block 1B and Block 2 –will lift 34-45 t to TLI. This paper will discuss the status of testing and integration for the Artemis I vehicle, manufacturing progress for the second vehicle and the manifest outlook for primary, co-manifested and secondary payloads in the current deep space exploration environment.

Creech, Stephen D.↗
NASA NTRS2019

Human Mars Entry, Descent and Landing Architecture Study: Deployable Decelerators

NASA’s Entry, Descent and Landing Architecture Study uses a trajectory simulation framework to evaluate various technologies and concepts of operations for human scale EDL at Mars. The study results inform agency technology investments. This paper summarizes the design assumptions and analysis of two deployable entry concepts performed in Phase 2 of the study. The entry concepts include a rigid deployable called the Adaptable Deployable Entry Placement Technology and an inflatable concept called the Hypersonic Inflatable Aerodynamic Decelerator. This paper describes the concept operations of these vehicles to deliver a 20-metric ton payload to the surface of Mars. Details of vehicle design and flight performance are summarized along with results of analysis on the aft body heating and its effect on the payload. Finally, recommended technology investments based on the results are presented.

Dwyer Cianciolo, Alicia↗
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 NTRS1993

A hypersonic vehicle approach to planetary exploration

An enhanced Mars network class mission using a lifting hypersonic entry vehicle is proposed. The basic vehicle, derived from a mature hypersonic flight system called SWERVE, offers several advantages over more conventional low L/D or ballistic entry systems. The proposed vehicle has greatly improved lateral and cross range capability (e.g., it is capable of reaching the polar regions during less than optimal mission opportunities), is not limited to surface target areas of low elevation, and is less susceptible to problems caused by Martian dust storms. Further, the integrated vehicle has attractive deployment features and allows for a much improved evolutionary path to larger vehicles with greater science capability. Analysis of the vehicle is aided by the development of a Mars Hypersonic Flight Simulator from which flight trajectories are obtained. Atmospheric entry performance of the baseline vehicle is improved by a deceleration skirt and transpiration cooling system which significantly reduce TPS (Thermal Protection System) and flight battery mass. The use of the vehicle is also attractive in that the maturity of the flight systems make it cost-competitive with the development of a conventional low L/D entry system. Finally, the potential application of similar vehicles to other planetary missions is discussed.

Murbach, Marcus S.↗
NASA NTRSDate not supplied

VERNE: Revealing the Mysteries and Histories of Venus

Introduction: The three Venus missions that were recently selected for upcoming flight (VERITAS, DAVINCI+, and EnVision) will be incredibly valuable to our understanding of Venus’ history, geology, and atmosphere. However, even once completed, key gaps in our knowledge of Venus, and more generally the formation and active processes on rocky, Earth-like planets, will still persist. Remaining questions include 1) how global intrinsic magnetic fields might be maintained on rocky worlds, and how they could then go extinct, and 2) what role atmospheric sulfur chemistry plays in climates of Earth-like planets, which is an increasingly timely subject as Earth’s own atmospheric sulfur content is climbing due to human activity. These questions require in-situ observations from Venus’ cloud deck, at the altitudes at which the UV absorber exists. The Venus Environment Research and Novel Exploration (VERNE) mission will address these questions with an aerial platform that will drift around the equatorial region of the planet for 9 days. VERNE will collect data to determine the identity of the mysterious UV absorber, while also taking magnetic field measurements over the tesserae, the regions on Venus that are most likely to retain remanent crustal magnetization, in order to understand the potential role of a past intrinsically-generated global magnetic field on Venus. Mission Objectives: The two major science objectives that drive the VERNE mission are: 1) Determine the identity of the Venusian unknown ultraviolet absorber(s). First observed approximately a century ago [1], the composition of Venus’ ultraviolet (UV) absorber is one of the oldest mysteries in Venus atmospheric chemistry [2,3]. While several candidate UV absorbers (mostly sulfur species) have been proposed, no consensus has been reached on its composition and its specific interactions with the atmosphere. Determining the identity of the UV absorber will aid climate models by showing how and where incident solar energy is absorbed by the atmosphere and make progress towards understanding the chemical and energetic processes taking place above Venus’ upper cloud deck [4]. 2) Determine if Venus retains evidence of a past, internally-generated magnetic field. While Venus does not currently have an intrinsically-generated magnetic field, evidence for the existence of a past field on Venus and a timeline of its decay will fill in a more holistic picture of the evolution of Venus’ geological record and atmosphere. As the oldest geologic units on the surface, Venus’ tesserae may still have remanent crustal magnetization signatures within the rocky composition [5]. In any case, the signatures detected will provide insight into Venus’ past geological and core dynamo activity. Mission Summary: VERNE includes a 3-part flight system made up of 1) an entry system with a HEEET (Heatshield for Extreme Entry Environment Technology) aeroshell, 2) an orbiter for data relay, and 3) an in-situ balloon and gondola. After entry into Venus’ atmosphere, the balloon will be deployed within the upper cloud deck, at an altitude of 62 km, over a tesserae region. The in-situ data collection will last for the duration of 2 full circumnavigations of the planet, which will take ~9 days. Instrumentation Suite: The four instruments that comprise VERNE’s instrument payload will enable the identification of the unknown UV absorber and the detection of remanent crustal magnetization if it exists in the tesserae. The proposed instrument suite cycle during mission operations is shown in Fig. 1. The four instruments are described below: 1) Adams (ion neutral mass spectrometer) has a range of 18-257 AMU and will detect and distinguish the mixing ratios of O₂, H₂O, H₂SO₄, S, S₂, S₃, S₄, S₅, S₆, S₇, S₈, SO, SO₂, OSSO, SO₃, Cl₂, FeCl₃ and other trace sulfur and organic species. The spatial (longitudinal) and temporal (day/night) variations will be observed throughout 2 circumnavigations with a sample cadence of 12 minutes. 2) Shelley (nephelometer) will determine the size distribution of the aerosols (0.4 to 36 um) in the atmosphere. With a size resolution of <0.7 um, it can determine which mode of H2SO4 is present and characterize the large (>30 um) organic particles previously detected by the Venera and Galileo missions [6,7]. 3) Herbert (UV imager) will measure UV radiance at 283 nm (the wavelength of SO2 absorption) and 365 nm (the unknown part of the absorber). UV images will be taken concurrently with the INMS and nephelometer to correlate UV absorption with abundances of the UV absorber species. 4) Vonnegut (magnetometer) has a range of >600 nT and a precision and accuracy of 1 nT. If magnetized by a past field, the crust may have retained a magnetization of up to 3 A/m2 [5]. With a noise floor of 10 nT, the magnetometer will be able to detect RCM from an altitude of 62 km, even if the thickness of the magnetized crust is just 1 km (Fig. 2). Mission Concept Design: VERNE will be launched with a mass of 3300 kg in an intermediate-high performance class vehicle with a 4-m fairing. The 475-day mission includes 466 days for the cruise, coasting, and orbit initialization phases before entry, descent, and balloon deployment. During the 9-day science phase, the aerial platform will make 2 circumnavigations of the planet at an altitude of 62 km. The INMS and the nephelometer will acquire data for 2 hours during the daytime and nighttime during each circumnavigation, while the UV imager will be on for the duration of the daytime, and the magnetometer will be operational throughout the entirety of the science phase. Data will be stored and processed with the JPL-designed Sphinx command and data handling system. Data will be sent from the balloon to the orbiter using an S-band relay link, stored on the orbiter, and then forwarded to Earth where it will be received by the DSN. Conclusion: VERNE will fill key gaps in our understanding of the history and ongoing processes related to the geology and atmosphere of Venus and rocky worlds in general. Even with adequate flight system contingencies and expected costs below the $900M New Frontiers cost cap, VERNE is not without its risks and challenges. Further trade spaces to explore include 1) using solely battery power vs. including solar panels to increase the mission duration and 2) investigating the use of lightweight materials and 3D-printed structures to reduce the gondola mass, among others. Acknowledgments: We would like to thank the JPL Planetary Science Summer School, especially our mentors Troy Hudson, Karl Mitchell, and Leslie Lowes, as well as our Team-X study lead Al Nash and the members of Team-X. We’d additionally like to thank our review panel for asking insightful questions and providing valuable feedback. References: [1] Ross, F. E. (1928) Astrophysical J., 68, 57-92. [2] Rossow, W. B. et al. (1980) J. Geophysical Research, 85, 8107-8128. [3] Pinto, J. P. et al. (2021) Nature Communications, 12, 175. [4] Titov, D. V. et al. (2007) Cosmic Research, 21, 401. [5] O’Rourke, J. et al. (2019) Geophysical Research Lett., 46, 5768–5777. [6] Limaye, S. S. et al. (2018) Astrobiology, 18(9), 1181-1198. [7] Grinspoon, D. H. et al. (2013) Planetary and Space Sci., 41(7), 515-542. [8] Parker, R. L. (2003) J. Geophysical Research, 108, 5006. *The cost information contained in this document is of a budgetary and planning nature and is intended for informational purposes only. It does not constitute a commitment on the part of JPL and/or Caltech.

H Alpert↗