IPPW-13 Program Abstracts - Modeling, Simulation & Testing...

Modeling, Simulation & Testing Session Thursday, June 16, 2016 from 8:30 AM to 11:45 AM

Conveners: Aaron Stehura Doug Adams Jason Rabinovitch Mars 2020 Entry, Descent and Landing Atmosphere Characterization

Gregorio Villar [email protected] 1

Mars 2020 Entry, Descent, and Landing Modeling and Performance

Soumyo Dutta [email protected] 2

EXOMARS 2016 LANDING IMPACT VERIFICATION F. Del Campo [email protected]

3

Impacting Earth: Testing and Analysis for Mars 2020 and Future Missions

Scott Perino, Louis Giersch, Velibor Cormarkovic, Zach Ousnamer, Darren Cooper, Brian Lim, Mike Johnson, and Gregory Peters

[email protected] 4

Mission-Level Requirements Validation for Atmospheric Entry Descent Systems

R. T. Stevens, K. E. Hibbard, D. S. Adams, J. M. O’neil, D. M. Gers, A. J. Shishineh, D. E. King, R. P. Roger, K. M. Siegrist, J. C. Taylor, and J. S. Graham

[email protected] 5

A SYSTEMS APROACH TO PLANETARY PROTECTION FOR BEE: A BIOSIGNATURE EXPLORER TO SAMPLE PLUMES OF EUROPA

Peter Spidaliere [email protected]

6

Modeling the Exo-Brake And The Development Of Strategies For De-orbit Drag Modulation

Marcus S. Murbach [email protected]

7

Kentucky Re-entry Universal Payload System Justin M. Cooper, Joseph K. Stieha, Alex M. Fowler, Alexander Martin

[email protected] 8

UAVs FOR TITAN EXPLORATION: FROM DESIGN TO POSTFLIGHT OF EARTH SCALED TEST VEHICLES.

Davide Bonetti [email protected]

9

NEW METHOD OF PREDICTING REENTRY RADIATION ON THE PRESENCE OF A NON-CONVENTIONAL/NON-CHARACTERIZED THERMAL PROTECTION SYSTEM MATERIAL

Gilles BAILET [email protected] 10

An Analytical Hypersonic Boundary Layer Transition (HBLT) Prediction Tool for the Sharp Edge Flight Experiment (SHEFEX) -III Reentry Vehicle

Siddharth Krishnamoorthy, Peter Noeding [email protected]

11

Hypersonic Entry Trajectory Optimization of High Ballistic Coefficient Vehicles

Christopher Lorenz [email protected] 12

Rocket Deployed Stability Flight Testing for a Candidate Aeroshell Shape

Alan L Strahan [email protected] 13

Dynamic stability of ERC via balloon drop tests Steve Lingard [email protected]

14

Dynamic CFD Simulations of the Supersonic Inflatable Aerodynamic Decelerator (SIAD-R) Ballistic Range Tests

Joseph Brock [email protected] 15

Preliminary Results of a Ballistic Range Test of a 2% Scale Mars 2020 Entry Capsule with Onboard Pressure Measurement

Mark Schoenenberger [email protected]

16

TBD

IPPW-13 Program Abstracts - Modeling, Simulation & Testing Session

July 13-17, 2016 Laurel MD USA http://ippw2016.jhuapl.edu/ 1

Mars 2020 Entry, Descent and Landing Atmosphere Characterization. Gregorio Villar1, Allen Chen1, Mi-chael Mischna1, David Kass1, Soumyo Dutta2, David Way2, Daniel Tyler3, Jeff Barnes3, Scott Rafkin4, Jorge Pla-Garcia4, 1Jet Propulsion Laboratory, California Institute of Technology, 2NASA Langley Research Center, 3Oregon State University, 4Southwest Research Institute.

Abstract – The Mars 2020 (M2020) Council of

Atmospheres (CoA) is a joint engineering and science team that is tasked with assessing atmospheric risk associated with entry, descent and landing (EDL). This paper presents the teams, tools, and processes involved in generating the atmospheric data that are used in EDL performance simulations. The overall methodol-ogy used by the M2020 CoA is largely the same as the Mars Science Laboratory CoA [1]. Mars Mesoscale Model 5 (MMM5) at Oregon State University and Mars Regional Atmospheric Modeling System (MRAMS) at the Southwest Research Institute are mesoscale models that generate atmospheric parame-ters, such as wind and density profiles, at the candidate landing sites and at the time of M2020 EDL. Prelimi-nary analysis shows that atmospheric conditions at the candidate landing sites do not significantly affect EDL performance. In fact, simulating EDL with mesoscale winds, instead of generic engineering winds, produces smaller landing ellipses. The M2020 CoA is preparing for the third landing site workshop in January 2017 by evaluating the candidate landing sites at nominal at-mospheric conditions, assessing the affects of dust events on EDL performance, and tuning the mesoscale models as more data is received.

References: [1] Ashwin Vasavada. et al. Assessment of Envi-

ronments for Mars Science Laboratory Entry, Descent, and Surface Operations, Space Science Review (2012) 170:793–835.



MARS 2020 ENTRY, DESCENT, AND LANDING MODELING AND PERFORMANCE. S. Dutta1, D.W. Way2, A. Chen3, and P. Brugarolas4. 1NASA Langley Research Center (Hampton, VA 23681, [email protected]), 2NASA Langley Research Center (Hampton, VA 23681, [email protected]), 3Jet Propulsion Laboratory (Pasadena, CA, 91009, [email protected]), and 4Jet Propulsion Laboratory (Pasadena, CA, 91009, [email protected]).

Abstract: Mars 2020 is the next planned U.S. rov-

er mission to land on Mars based on the design of the successful 2012 Mars Science Laboratory (MSL) mis-sion [1]. It will launch in July 2020 and will reach Mars in February 2021. Mars 2020 retains most of the entry, descent, and landing (EDL) sequences of MSL, including the closed-loop entry guidance scheme based on the Apollo guidance algorithm and the Skycrane landing system [2]. However, a few improvements to the EDL system design have been made during the design process. The Mars 2020 will trigger the para-chute deployment and descent sequence using a range trigger instead of a velocity trigger. This change will vastly improve the landing precision of the vehicle. Additionally, Terrain Relative Navigation (TRN) has been incorporated into the system, allowing the vehicle to safely land at a more hazardous landing site than was possible for MSL. These and other changes have led to an effort to improve the modeling and simulation environments used for MSL and modify them to suit the needs of the Mars 2020 mission.

This paper will focus on the models and simula-tions developed to predict EDL performance for MSL [3, 4], how these models have been updated to charac-terize the Mars 2020 entry environments, and finally show EDL performance results generated from the simulation. Initially, improvements were made to recti-fy differences seen between MSL’s pre-landing simu-lation predictions and as-flown performance. These differences included improving the knowledge of the local gravity field in the navigation flight software and simulation, since underestimating this value led to lower than expected touchdown velocities for MSL [3]. Additionally, there has been a need for the Mars 2020 mission to characterize EDL performance at spe-cific landing sites earlier than it was done during MSL’s design process. Improved landing hazard maps and mesoscale atmospheric tables are used within the simulation to generate predictions specific to a landing site. The enhanced modeling and simulation tool has allowed the design team to predict the EDL perfor-mance at eight potential landing sites and has aided the work of project scientists and the Mars 2020 landing site workshop participants as they downselect landing sites for the mission.

References: [1] Mustard, J., et al. (2013) “Report of the Mars

2020 Science Definition Team,” Tech. rep., Mars Ex-ploration Program Analysis Group (MEPAG). [2] Steltzner, A. (2013) “Mars Science Laboratory Entry, Descent, and Landing System Overview”, AAS 13-236. [3] Way, D. W., Davis, J. L, and Shidner, J. D. (2013) “Assessment of the Mars Science Laboratory Entry, Descent, and Landing Simulation”, AAS 13-420. [4] Way, D. W. (2013) “Preliminary Assessment of the Mars Science Laboratory Entry, Descent, and Landing Simulation”, IEEE-2013-2755.



EXOMARS 2016 LANDING IMPACT VERIFICATION F. del Campo1, Y. E. Jauregui2 and R. Haya3, 1SENER Ingeniería y Sistemas Avda. Zugazarte, 56-48930 Las Arenas – Spain [email protected], 2SENER Ingeniería y Sistemas Avda. Zugazarte, 56-48930 Las Arenas – Spain [email protected], 3SENER Ingeniería y Sistemas Severo Ochoa, 4 (PTM) - 28760 Tres Cantos – Spain [email protected],

Introduction: The ESA mission ExoMars 2016

has an Entry, Descent and Landing Demonstrator Module (EDM), named 'Schiaparelli', including a Sur-face Platform (ESP) which will land on Mars by means of a Crushable Structure absorbing the final impact energy. The spacecraft carrying the Module was launched from Baikonur, Kazajstan, on 2016.03.14 and is on its way to Mars, where the Module will be landed on 2016.10.19. SENER, with TAS-I as Prime Contrac-tor, is responsible for the Crushable Structure: a low-density deformable honeycomb panel, designed to pro-tect the scientific payload at impact event.This paper is focused on the dynamic analyses and tests performed to demonstrate the reliability of the ESP touchdown.

Figure 1 ExoMars 2016 Crushable Structure ready

for a landing test Landing System evolution: The use of a crushable

structure as main landing system for a planetary probe is not new. It was for instance used in the landings of Mars 2, 3 and 6. In Exomars it has been adopted as a result of the development of the program. Thus, the mission configuration has evolved from the single mis-sion carrying the rover using hyperbolic entry and 1 Ton at entry using Sozuz as launcher (2005) up to the actual split into two missions, first with NASA and then with Roskosmos, in which EDL for the recently launched mission became a demonstrator weighting 600 kg at entry. The paper presents a summary of the mission evolution leading to the adoption of the crush-able structure as landing system in order to meet the requirements for the actual mission configuration.

Verification strategy: One of the major challenges of any space mission involving planetary Landers is to predict realistic impact conditions, in terms of impact loads and stability, considering very wide uncertainty ranges at touchdown. The main novelty introduced in this paper is the presentation of how this problem has been faced for ExoMars 2016.

Statistical Approach: The adopted solution is based on a Statistical Approach in which the impact loads are derived statistically from a set of random landing cases. In this approach, all relevant conditions and parameters are implemented in MSC.Dytran nonlinear simulations as random variables; i.e. landing velocity and attitude, ambient temperature, terrain slope and properties and rock layout. This original approach is in stark contrast to the deterministic methods considered for other plan-etary missions.

Soil Model: A second novelty is the study of the in-teraction between the Lander and the Martian soil. Un-like other missions in which the soil properties are not considered in the analysis, a soil mathematical model has been developed and implemented. In fact, the ge-otechnical properties of the ExoMars 2016 landing site (Meridiani-Planum) are derived from tests performed with Mars soil simulants. The soil model is based on a Drucker-Prager yield surface combined with a hydro-static pressure - volumetric strain function.

Testing Campaign: The simulations are validated

thanks to a test campaign in which a full-scale Qualifi-cation Model of the ESP (with a flight-model repre-sentative Crushable Structure) is tested under different extreme landing conditions; including various rock sizes, terrain slopes and impact velocities. In addition to this, some tests are performed in Mars like pressure (around 10 mbar). Finally, the paper presents the corre-lation between the mathematical model and the real tests, providing evidence of how the analytical model represents reality with a high degree of similarity.



mailto:[email protected]



Lessons Learned: Lessons learned from analysis,

tests and correlation, useful for other present and future Space Projects, will be provided in the paper.

Other Applications: The successful qualification of the Exomars’16 crushable structure in which the predictability enabled by current computing capability fosters its application as landing system for medium-small size probes. An initial outlook of potential appli-cations beyond Exomars will be presented, covering planetary and earth return probes.



Impacting Earth: Testing and Analysis for Mars 2020 and Future Missions Scott Perino1†, Louis Giersch1, Ve-

libor Cormarkovic1, Zach Ousnamer1, Darren Cooper1, Brian Lim1, Mike Johnson1, and Gregory Peters1

1Jet Propulsion Laboratory, California Institute of Technology, 4800 Oak Grove Drive, Pasadena, CA 91109, †[email protected]

Disclaimer: Although NASA has no official plans

at this time for a mission to return samples from Mars,

the Program Formulation Office of the Mars Explora-

tion Program sponsors ongoing mission concept studies,

systems analyses, and technology investments which

explore different strategies for the potential return of

samples from Mars, consistent with the charter of the

program and stated priorities of the science commu-

nity.

Background: The proposed Mars Sample Return

(MSR) campaign [1] architecture calls for three

launches from Earth, spaced over several years. The

first mission, called Mars 2020, is a rover that collects

and preserves Martian geological samples inside spe-

cially designed tubes. The next mission called Sample

Return Lander (SRL), is a rover with an attached launch

vehicle used first, for collection and then, for Mars orbit

insertion of the samples within in a spherical container

called an Orbiting Sample (OS). The last mission called

Sample Return Orbiter (SRO), is a spacecraft that pro-

vides communications relay, orbital rendezvous and

capture of the OS, and Earth-return delivery of the sam-

ple container. After the SRO captures the OS it is then

inserted and sealed within an onboard rentry vehicle

called an Earth Entry Vehicle (EEV) [2]. Once the SRO

approaches Earth, the EEV is released from the SRO,

reenters the Earth’s atmosphere, and intentionally im-

pact lands without parachute on a designated landing

site at the Utah Training and Testing Range (UTTR) [3].

For the MSR campaign to be a success, the Mars

2020 mission must know, and design for, the require-

ments associated with the as-yet unfunded subsequent

MSR missions. To address this issue, JPL has assem-

bled the Mars Advanced Development Team (ADT).

Acting on behalf of the downstream missions, ADT per-

forms testing, analysis, and design explorations, which

then inform requirements that are levied onto Mars

2020.

Abstract: The culminating event of the MSR cam-

paign occurs when the EEV impact lands at UTTR and

the martian samples are recovered for scientific analy-

sis. The impact landing is a design driver for many com-

ponents, specifically the sample tubes, the OS, and the

EEV. This paper describes ADT’s coordinated effort to

predict Earth impact loads and to develop preliminary

designs for the OS and other future MSR mission hard-

ware. The effort is focused into three main thrust areas:

1) Full scale soil impact tests, 2) High fidelity and

ADT thrust areas for Earth impact loads

predictive dynamic analysis, and 3) Preliminary hard-

ware development.

ADT and JPL’s Planetary Landing Testbed (PLT)

team, have constructed an 86 ft tall aluminium truss

frame tower for conducting impact tests into soft soil.

The tower has a soft soil bed at its base and is fitted with

a pully system connected to a pneumatic piston and

pressure tank. The system can accelerate EEV-like pen-

etrometers and other hardware into the soil at speeds in

excess of the expected terminal velocity of an EEV. To

date, 15 impact tests have been conducted using the sys-

tem. Sphere-cone penetrometers ranging in mass from

40 kg - 140 kg have been impacted at velocities from 20

m/s - 67 m/s into soil with 5% - 31% water content. A

description of the impact tower and a video of an impact

event are presented.

Dynamic analysis closes the gap between impact

testing and preliminary hardware design. The analysis

focuses both on accurate modeling of soil as well as on

predicting the response of sample tubes and the OS im-

pacting against the soil. Peak acceleration results from

penetrometer impact tests have been validated to within

±20% across the soil moisture and velocity range. Fur-

thermore, numerous OS designs have been evaluated

and refined via dynamic analysis. Results from the soil

model validation and OS and tube analysis are pre-

sented.

Preliminary hardware design efforts were conduced

to ensure that the Mars 2020 tubes have the necessary

design features and load capacity to withstand all of the

downstream mission events, most notably Earth impact.

Several OS designs were created to determine how

tubes are inserted and secured within the it. Each design

was vetted through dynamic analysis and the most

rescent OS designs and how they interface with the

Mars 2020 tubes are presented.



References:

[1] Mattingly, R., and May, L. (2011) “Mars Sample

Return as a Campaign,” IEEE Aerospace Conference, Big

Sky, MT: IEEE, 2011, pp. 1–13

[2] Dillman, R., and Corliss, J. (2008) “Overview of the

Mars Sample Return Earth Entry Vehicle,” Sixth International

Planetary Probe Workshop, Hampton, VA: pp. 1–8.

[3] Fasanella, E., Jones Y., Knight, N., and Kellas, S.

(2001), “Low Velocity Earth Penetration Test and Analysis”

AIAA, A1338



Mission-Level Requirements Validation for Atmospheric Entry Systems

R. T. Stevens, K. M. Siegrist, J. M. O’Neil, D. M. Gers, A. J. Shishineh, R. P. Rogér, J. C. Taylor, J. S. Graham, D. E. King, D. S. Adams1, and K. E. Hibbard2

The Johns Hopkins University Applied Physics Laboratory 11100 Johns Hopkins Road

Laurel, MD 20723 [email protected], [email protected]

The Johns Hopkins University Applied Physics Laboratory (APL) is heavily engaged in all aspects of hypersonic flight and associated phenomenology through a multitude of programs involving aerodynamics, guidance and control, aerothermal heating, and ablation. This broad spectrum of programs involves widely varying regimes of hypersonic flight. For example, APL expertise in hypersonic reentry has supported the Department of Energy (DoE) with nuclear safety assessments of NASA missions including Ulysses, Galileo, Cassini, and New Horizons [1],[2],[3], all probes that use lightweight radioisotope heater units (LWRHUs). Characterization of reentry and impact conditions played a fundamental role in the analyses and testing undertaken to predict response of an LWRHU during reentry or impact. The most challenging reentry environment studied to date by APL was that of the Cassini mission, which swung-by Earth as part of its Venus-Venus-Earth-Jupiter gravity assist to reach Saturn. The APL-led risk study analyzed the LWRHU response in the event the probe were to reenter during the Earth swing-by, subjecting the LWRHU to the aerothermal hypersonic heating from the Mach 75 flow of a 19.5 km/s reentry. This case was potentially the highest energy level ever encountered by a manmade object during an Earth reentry. Less challenging thermal engineering scenarios for endo- and exoatmospheric ballistic missiles are routinely treated via long-established software implementations that continue to evolve to meet emerging challenges.

In recent years, interest has grown in obtaining physical samples of other bodies in the solar system, whether for sample return or for in-situ analysis. Such missions have driven a need for improved under-standing of Entry, Descent, and Landing (EDL) phases of space probe transit through atmospheres of Earth and other bodies in the solar system. Recognizing the growth in missions involving EDL, APL has undertaken a study to investigate how existing in-house capabilities and areas of expertise that were independently developed can be leveraged for application to planetary entries. The principal focus of this study is on the ability to perform high-level mission trade studies to help define mission-level requirements, with detailed analyses being conducted

at NASA’s Ames and Langley Research Centers. There is also an interest in exploring as-yet untapped analytical capabilities that have not yet been applied to planetary explorations. This paper will document some relevant past projects in aerodynamics, computational fluid dynamics (CFD), radiation transport, and ablation. A validation effort was undertaken using the Langley Aerothermodynamic Upwind Relaxation Algorithm (LAURA) [4],[5] and the Unstructured Three-Dimensional (US3D) [6],[7] CFD codes by comparing them with FIRE II reentry convective heating results [8]. The High-temperature Aerothermodynamic Radiation (HARA) [9],[10] code was used to calculate the radiation heating and then compared with FIRE II results [11],[12]. Ablation was modeled with the Charring Material Ablation (CMA) code [13] and then compared with results from the Stardust mission.

APL created an automated engineering tool to aid mission planning that could automate advanced design EDL analyses. It uses the Trajectory Analysis and Optimization System (TAOS) code [14] to calculate three- and six-degree-of-freedom trajectories. The heatshield required thickness is solved for using the CMA code. Three distinct analysis capabilities will be explained and demonstrated: (1) trajectory and Thermal Protection System (TPS) thickness optimization given a user defined metric and either initial entry conditions or impact location, (2) a Monte Carlo study of the trajectory and TPS response, and (3) a sensitivity study of the parameters used in the Monte Carlo study.

Lastly, Large-Eddy Simulation (LES) CFD solutions are shown for a generic Multi-Mission Earth Entry Vehicle originally conceived for the Mars Sample Return mission. The solutions come from an efficient, pseudo-spectral, high-order LES solver using Forza, an immersed boundary method developed at APL, and is a descendent of the well-validated JHU LES code [15]. A study is presented of unsteady subsonic and supersonic flow including the wake region of the reentry vehicle to examine unsteady aerodynamic forces through the transonic pinch point.



References

[1] E. F. Lucero et al., “Reentry Response of the Lightweight Radioisotope Heater Unit Resulting from a Cassini Venus-Venus-Earth-Jupiter Gravity Assist Maneuver Accident,” JHU/APL ANSP-M-24, DoE/NE/32170-24, (1996).

[2] Y. Chang et al., “New Horizons Launch Contingency Effort,” Space Technology and Applications Int. Forum, AIP Conf. Proc. 880, pp. 590–596 (2007).

[3] Y. Chang, “Aerospace Nuclear Safety at APL: 1997–2006,” Johns Hopkins APL Technical Digest, Vol. 27, No. 3, pp. 253–260 (2007).

[4] P. A. Gnoffo, R. N. Gupta, and J. L. Shinn, “Conservation Equations and Physical Models for Hypersonic Air Flows in Thermal and Chemical Nonequilibrium,” NASA TP-2867 (February 1989).

[5] P. A. Gnoffo, “An Upwind-Biased, Point-Implicit Relaxation Algorithm for Viscous, Compressible Perfect-Gas Flows,” NASA TP-2953 (1990).

[6] I. Nompelis, T. W. Drayna, and G. V. Candler, “Development of a Hybrid Unstructured Implicit Solver for the Simulation of Reacting Flows Over Complex Geometries,” AIAA 2004-2227 (June 2004).

[7] I. Nompelis, T. W. Drayna, and G. V. Candler, “A Parallel Unstructured Implicit Solver for Hypersonic Reacting Flow Simulation,” AIAA Paper 2005-4867 (June 2005).

[8] Hash et al., “FIRE II Calculations for Hypersonic Nonequilibrium Aerothermodynamics Code Verification: DPLR, LAURA, and US3D,” AIAA 2007-605 (January 2007).

[9] C. O. Johnston, B. R. Hollis, and K. Sutton, “Non-Boltzmann Modeling for Air Shock-Layers at Lunar Return Conditions,” Journal of Spacecraft and Rockets, Vol. 45, No. 5, pp. 879–890, (2008).

[10] C. O. Johnston, “Nonequilibrium Shock-Layer Radiative Heating for Earth and Titan Entry,” Ph.D. Dissertation, Virginia Polytechnic Inst. and State Univ., Blacksburg, VA (November 2006).

[11] C. O. Johnston, B. R. Hollis, and K. Sutton, “Nonequilibrium Stagnation-Line Radiative Heating for Fire II,” Journal of Spacecraft and Rockets, Vol. 45, No. 6, pp. 1185–1194 (2008).

[12] P. A. Gnoffo, C. O. Johnston, and R. A. Thompson, “Implementation of Radiation, Ablation, and Free-Energy Minimization in Hypersonic Simulations,” Journal of Spacecraft & Rockets, Vol. 47, No. 2, pp. 481-491 (2010).

[13] “User’s Manual Non-Proprietary Aerotherm Charring Material Thermal Response and Ablation Program CMA87S,” Acurex UM-87-13/ATD, 9187.

[14] D. Salguero, “Trajectory Analysis and Optimization System (TAOS) User’s Manual,” SAND95-1652 (1995).

[15] F. Portè-Agel, C. Meneveau, and M. B. Parlange, “A scale-dependent dynamic model for large-eddy simulation: application to a neutral atmospheric boundary layer,” J. Fluid Mech., Vol. 415, pp. 261-284 (2000).



A SYSTEMS APROACH TO PLANETARY

PROTECTION FOR BEE: A BIOSIGNATURE

EXPLORER TO SAMPLE PLUMES OF

EUROPA. T. Errigo1, M. Casey1, J. Garvin1, P. Spi-

daliere1, 1NASA Goddard Space Flight Center, Code

690, Greenbelt, MD 20771 (Therese.Errigo-

[email protected]).

Introduction: Icy bodies, and particularly those

with liquid oceans, are among the most likely places to

detect evidence of past or current non-terrestrial, mi-

crobial life. Accordingly, the Planetary Protection

(PP) requirements for these bodies are among the most

stringent possible. The probability of contaminating an

ocean or other liquid body must be less than 1 x 10-4

per mission.

We herein describe the integrated systems approach

taken to develop creative solutions for meeting this

requirement on the Biosphere Explorer for Europa

(BEE). This probe was designed by Goddard Space

Flight Center (GSFC) in response to a directive from

NASA to explore options for possible augmentations to

the Europa Multi Flyby Mission (EMFM) to directly

search for life at Europa. BEE was designed to be re-

leased from the EMFM, to actively target and fly

through a plume at a nominal altitude of 5 km, collect

sample plume material, and then search for molecular

signatures of life in these gas and icy particles vented

into space from Europa’s interior ocean.

Driving Concerns: Given the low altitude pass re-

quired to collect a sample and the possibility of a colli-

sion during the flythrough, the potential for contamina-

tion of Europa had to be carefully addressed. Like-

wise, without a divert maneuver after the flythrough, or

as the result of a failed divert maneuver, eventual colli-

sion with Europa was also a concern. In addition to

meeting planetary protection requirements and those

requirements levied by the BEE science goals, BEE

had to be designed to minimize technical, operational,

and programmatic impacts to the EMFM – goals and

requirements with a high likelihood of conflict, if not

managed properly.

The initial first cut attempt to meet planetary pro-

tection requirements by the BEE design team was

based on avoiding collision with the moon on the first

or subsequent orbits at all costs. Meeting this per-

ceived requirement would have required a level of reli-

ability that would have exceeded the core science mis-

sion needs (redundant, heavier, more expensive, and

with less science instrumentation to counterbalance

cost and mass impacts) and would have limited opera-

tional flexibility by forcing a release trajectory that

insured collision with another moon. In order to limit

recontamination of the hardware after bioburden reduc-

tion, we initially designed a flight worthy and deploya-

ble, bio-barrier (a “suitcase” or “coffin”) that would

protect the BEE until after launch. However, upon

further study, we discovered other less burdensome

options. These other approaches would not only result

in a lower cost, but also a better design that would

maximize science return and minimize impacts to the

EMFM, while meeting vitally important Planetary Pro-

tection requirements with significant margin.

BEE Planetary Protection Stategy: The BEE

Planetary Protection approach was built primarily upon

a thorough bioburden reduction strategy and the use of

radiation to kill any microorganism on exterior surfaces

that could be contaminated at launch or by EMFM dur-

ing flight.

The latter approach, use of radiation, incorporated

methods to build-up dose on the exterior surfaces by:

(1) designing an outer skin surface that would act both

as a biobarrier and as a simplistic geometry that would

prevent significant shielding of microorganisms from

radiation (nooks and crannies), (2) avoiding collision

with the moon until sufficient dose accumulated and

(3) demonstrating that even if a collision were to occur

that the hypervelocity impact would leave no debris

large enough such that it would shield the microor-

gasims from the intense Jovian radiation environment.

The hypervelocity analysis demonstrated that the exte-

rior thin skin cover would vaporize upon impact there-

by killing most microorganisms. Any remaining space-

craft debris raining down on the surface would be too

small to provide effective shielding. The resultant

probability of delivering microorganisms to the liquid

ocean beneath Europa’s thick ice, even over long time

scales, would meet Planetary Protection requirements

for icy bodies.

A key component to this strategy was insuring that

the hardware under the skin would be treated to reduce

the bioburden to acceptable levels. BEE Bioburden

reduction strategy described herein leverages tech-

niques developed and used for GSFC’s Sample Analy-

sis on Mars (on the Mars Science Laboratory) and

MOMA (European Exomars) instruments. The BEE

Bioburden approach included: high temperature Dry

Heat Microbial Reduction (DHMR); specialized clean-

ing and aseptic assembly processes developed at

GSFC; a modular design of the spacecraft such that

DHMR could be performed at a high level of assembly

to reduce schedule and cost impacts; early considera-

tion of alignment issues raised by high temperature

DHMR; and an exterior skin that would protect most

components, acting both as a bioburden barrier for

surfaces underneath it and as a simple, smooth surface

that would support the radiation sterilization strategy.





Details of the BEE Planetary Protection approach

are included, along with predictions for dose accumula-

tion on the surface as a function of shielding and a con-

servative probabilistic assessment of the likelihood of

contamination of Europa’s ocean.

The approach decribed herein would minimize im-

pacts to EMFM, both technical and programmatic.

More importantly, it would meet the Agency’s Plane-

tary Protection requirement with a healthly margin and

would enable low altitude collection of plume material

before organics could precipitate out and where the

concentration of those molecules would be the greatest.



Modeling the Exo-Brake And The Development Of Strategies For De-orbit Drag Modulation M. S. Murbach,1 P. Papadopoulos,2 C. Glass, 3 A. Dwyer-Ciancolo, 3 R.W. Powell, 3 S. Dutta, R. Carlino,3 F. A. Tanner,1 R. Carlino, 1 1NASA Ames Research Center, Moffett Field, CA 94035, 2San Jose State University, Aero-nautical Engineering Department, One Washington Square, San Jose, CA, 95192, 3NASA Langley Research Center, 8 Lindburgh Way, Hampton, VA, 23681

Abstract: The Exo-Brake is a simple, non-propulsive means of de-orbiting small payloads from orbital plat-forms such as the International Space Station (ISS). Two de-orbiting experiments with fixed surface area Exo-Brakes have been successfully conducted in the last two years on the TechEdSat-3 and -4 nano-satellite missions. The development of the free molecular flow aerodynamic data-base is presented in terms of angle of attack, projected front surface area variation, and altitude. Altitudes are considered ranging from the 400km ISS jettison altitude to 90km. Trajectory tools are then used to predict de-orbit/entry corridors with the inclusion of the key atmospheric and geomagnetic uncertainties. Control system strategies are discussed which will be applied to the next two planned TechEdSat-5 and -6 nano-satellite missions – thus in-creasing the targeting accuracy at the Von Karman altitude through the proposed drag modulation tech-nique.



The Kentucky Re-entry Universal Payload System: Analysis and Preliminary Subsystems.

Justin M. Cooper, Joseph K. Stieha, Alex M. Fowler, and Alexandre Martin

Department of Mechanical Engineering, University of Kentucky, Lexington, KY

Introduction: The Kentucky Re-entry Universal

Payload System (KRUPS) is a low-cost, microsatellite

re-entry vehicle intended to collect real time data and

aid analysis and reconstruction of thermal protection

systems. Its primary mission is to serve as a tool for

researchers to analyze thermal response of re-entering

spacecraft. Because of its size and manufacturing tech-

nique, KRUPS is envisioned as a cheap launch alterna-

tive to large scale institutional spacecraft. This paper

presents a preliminary analysis of 1D thermal response

and CFD based on a ballistic trajectory and subsystem

design. In addition, an in-house 6-DOF trajectory code

designed to evaluate the boundary conditions used

which the CFD code is also presented.

Movitation: Current modeling techniques expose

various gaps in the EDL, mainly pointed towards flight

data reconstruction. The main problem revolves around

correctly evaluating the heat fluxes at the surface of the

vehicle. To facilitate this process, a statistically mean-

ingful sample of re-entry data is needed. To this end,

KRUPS provides a cheap and easily manufacturable

system that can transmit valuable information about the

entry process, such as isothermal contours and pres-

sure. To ensure that this data is meaningful to the entry

physics community, an analysis has been proposed for

a ballistic decay trajectory of said spacecraft to validate

the enthalpy regime of specific charring ablators of

interest. Additionally, the calculated trajectory and

subsequent analysis provides an estimate of the data

acquisition and transmission capabilities of the space-

craft.

Figure 1. KRUPS Cross Section

Analysis: The KRUPS Aerodynamic profile is

based on the Deep Space Mars Microprobe and Galileo

spacecraft architectures[1-2]. The trajectory code is

useds a 6-DOF system, and is based on the work of

Vinh et al.[3] Initial estimates of the flow field charac-

terization and the material response are done using a

University of Kentucky in-house codes that solves both

the Navier-Stokes equations and in-depth ablation[4].

All of the electronic systems are integrated open-source

hardware that allow for ease of use, and allows for fu-

ture users to design projects without knowledge in

electronics.

Figure 2. Data Acquisition System

References:

[1] Mitcheltree R. A., Moss J. N., Green F. A.

(1999) Journal of Spacecraft and Rockets, 36.3, 392–

398.

[2] Milos, F. S., Chen, Y. K., et. al. (1999) Journal

of Spacecraft and Rockets Journal of Spacecraft and

Rockets., 36(3), 298-306.

[3] Vinh N. X. (1980) Hypersonic and Planetary

entry flight mechanics, Technical Report A 81, 16245.

[4] Weng, H., Zhang, H., Martin, A. (2012) Jour-

nal of Thermophysics and Heat Transfer, 29(4).



UAVs FOR TITAN EXPLORATION: FROM DESIGN TO POSTFLIGHT OF EARTH SCALED TEST

VEHICLES.

D. Bonetti and A. Latorre,

DEIMOS Space S.L.U., Ronda de Poniente 19, Tres Cantos, 28760, Spain

([email protected], [email protected])

PERIGEO is a project funded by the INPPRONTA

programme (CDTI, Spain, 2011-2014) which aims to

provide a framework for research and validation of

space technology and science by means of Earth-

analogue environments and to apply this framework in

a number of specific technological developments.

Within PERIGEO, tools and methodologies are de-

signed and implemented to continue the efforts per-

formed in laboratory development, exposing the tech-

nology to representative and, to a certain extent, realis-

tic dynamic and environmental conditions, similar to

those of space missions.

Specific goals in PERIGEO consist in the follow-

ing: defining new missions and vehicles by multi-

disciplinary design optimization (MDO) of key charac-

teristics; developing a robust and failure-tolerant con-

trol system; creating an integrated research and design

process and tools to mature the project developments

through a logical work sequence (research, laboratory-

testing, real-flight-"analogue" testing) exploring new

navigation methods based on imaging and inertial sen-

sors, including its hybridization; and advancing in new

fault tolerance techniques for space computer architec-

tures.

Among different scenarios considered within the

project, this paper presents the activities covered in the

area of “Atmospheric Flight” tackling the goals of the

MDO and the research-design-test areas. More in de-

tails, the mission scenario considered for the applica-

tion of the design process and tools is that of UAVs

flying in Titan atmosphere, collecting environment data

to contribute to the in-situ scientific exploration of the

Saturn's moon. Two mission scenarios emerged as the

most interesting within PERIGEO:

1) The first concept is a Titan UAV glider mission.

The core exploration mission is performed by a lander

vehicle, placed on the surface of Titan by an EDL sys-

tem. The glider UAV separates from this principal

lander mission several km above ground and acts as a

scout pathfinder, exploring the surroundings from

above the lander. It takes 3D stereo pictures and sends

them to the lander making it possible to build a 3D

model of the surroundings to take better decisions on

the paths to be followed during its ground exploration.

2) The second concept is an innovative multi-UAV

mission where 5-6 UAVs fly in a formation to explore

as a team the Titan environment. The key benefit of

this approach is to split the different payloads initially

planned to be loaded onboard of a single vehicle

(AVIATR-approach) achieving in this way a much

lighter and flexible UAV system. All the UAVs share

the external shape, the mass and key subsystems (i.e.

power & propulsion) to reduce costs while specific

different payloads are loaded in each UAV according

to his role in this collaborative mission scenario.

Earth-scaled UAVs have been designed, built and

tested in wind tunnel and in flight under aerodynamic

equivalent conditions. Focusing on the glider concept

introduced, the paper presents the full process from

desing to flight of the Earth-scaled model and the re-

sults of the postflight analysis.

Analogue mission testing on Earth for atmospheric

flight missions has been demonstrated in the project,

with affordable cost thanks to an efficient methodology

[1]. Aspects requiring more detailed analysis for the

multi-UAV missions are also highlighted: this chal-

lenging scenario could be an interesting framework for

international collaborations in view of future Titan

exploration missions.

References:

[1] Bonetti D. et al (2015) “MDO, Wind Tunnel and

Flight Tests of UAVs for Titan exploration”, DASIA 2015.





NEW METHOD OF PREDICTING REENTRY RADIATION ON THE PRESENCE OF A NON-CONVENTIONAL/NON-CHARACTERIZED THERMAL PROTECTION SYSTEM MATERIAL. G. Bai-let1, A. Bourgoing2, T. Magin3 And C. O. Laux4, 1PhD Candidate, Ecole Centrale Paris, Laboratoire EM2C, CNRS UPR288, 92290 Châtenay-Malabry, France . [email protected], 2AIRBUS Defence & Space, Route de Verneuil, 78133 Les Mureaux CEDEX, France, 3Associate Professor, Von Karman Institute for fluid dynamics, Aeronautics & Aerospace Dept, Chaussée De Waterloo 72, 1640 Rhode-St-Genèse, Belgium, 4Proffesor, Ecole Centrale Paris, Laboratoire EM2C, CNRS UPR288, 92290 Châtenay-Malabry, France.

Introduction: Reentry radiation is one of the lead-

ing phenomena to take in account while designing a high speed reentry vehicle (>10km/s). In the case of an ablative Thermal Protective System (TPS), the physi-co-chemical processes involved are complex and rarely fully characterized.

The paper is presenting a novel method based on a mixed numerical/experimental approach allowing to predicting the radiative flux with the effect of ablation.

To illustrate this new method, the radiation of the QARMAN mission (Flight 2016 Q3) [1] will be pre-dicted. The reentry speed is 7.8km/s at 120km of alti-tude. This Low Earth Orbit (LEO) return mission is equipped with a cork based TPS material (Cork P50). This TPS material is widely used in the industry but the characterization is complex (natural phenols, wide variety of molecules) and high margins are used to reduce cost [2].

Brief description of the method: The idea behind the method is to assess the absorption and emission of a hot/ablative boundary layer when a TPS sample is subjected to a high enthalpy flow (ICP Torch).

By characterizing the radiation environment of the facility and comparing it with simulations (no ablation case) and experiments (with a spectrometer inside the TPS sample), it is possible to clearly predict the radia-tion of any mission with any TPS material with classi-cal contributions (CN, CO…) or more exotic ones (phenols, Alcalins, Metallic ions…).

Validation: Before predicting any mission cases, a validation with a basic and fully characterized material (Graphite) has been done. Nowadays, the numerical models are able to predict the radiation of a fully car-

bonaceous [3] TPS material such as pure Graphite. The validation case will permit to quantify the uncertainties of the method. Nevertheless, the numeri-cal/experimental approach will permit to observe arti-facts of human handling (Na+…) where the numerical codes will never be able to do so.

LEO type reentry prediction: The QARMAN vehicle (34cm long for 5kg) will reenter the Earth’s atmosphere from an initial altitude of 400km by the end of 2016. Within the payload bay, a emission spec-trometer will take radiations of the stagnation line. The paper will expose the results of applying (exemple in Fig. 1) the developed and validated new method to predict radiation in the presence of a non-characterized TPS material (Cork P50) to this reentry mission.

Conclusion: The new method is presented as an easy to use and widely usable technic to predict the radiation of a reentry mission on the presence of a non-conventional/non-characterized thermal protection system material. The success of the QARMAN mis-sion will permit to compare the predictions with flight data and permits to use this method to design other vehicles and perform fundamental research.

References: [1] I. Sakraker, et al., Qarman: an Atmospheric En-

try Experiment on CubeSat Platform, 8th ESASV, Por-tugal, March 2015.

[2] NASA-CR-170881, SRB TPS Materials Test Results in an Archeated Nitrogen Environment, Lock-heed Missiles and Space Co, 1979.

[3] A. Turchi, et al., On the Flight Extrapolation of Stagnation-Point Ablative Material Plasma Wind Tun-nel Tests, 8th ESASV, Portugal, March 2015.

Figure 1: Calculated spectrum (61km of altitude, 5,3km/s, no ablation and cold wall, up) and overlayed in red, the prediction of

QARMAN radiation taking in account the absorption/emission of the hot ablative boundary layer (down)

Ca II



AN ANALYTICAL HYPERSONIC BOUNDARY LAYER TRANSITION (HBLT)

PREDICTION TOOL FOR THE SHARP EDGE FLIGHT EXPERIMENT (SHEFEX) –III

REENTRY VEHICLE. S. Krishnamoorthy1 and P. Nöding2

1PhD. Candidate, Department of Aeronautics and Astronautics, Stanford University, Stanford, CA,

USA. Email address: [email protected] 2Thermal Engineering Department, Airbus Defence and Space, Bremen, Germany

Email address: [email protected]

Introduction: Hypersonic Boundary Layer

Transition (HBLT) has been studied for over fifty

years. These studies have been theoretical,

experimental and computational in nature.

However, the phenomenon of transition to

turbulence has proved to be a very complex one; no

effective way to predict boundary layer transition

has been found yet. This problem remains an

extremely important one to study, especially for

heating and flight dynamics considerations for

hypersonic vehicles [1]. Transition is affected by a

multitude of factors ranging from flow conditions

and noise environment to vehicle geometry and

trajectory. It is also difficult to isolate the effect of

any of these factors because of their complex

interaction with each other. Current transition

prediction methods range from purely empirical

correlations to complex, transport equation-based

turbulence models. While the empirical

correlations are fastest to compute, they are often

the least accurate and most problem-specific.

However, they provide a fast engineering estimate

of transition location for more detailed analyses.

The current work focuses on the creation and

verification of a hypersonic boundary layer

transition prediction tool for the SHarp Edge Flight

EXperiment (SHEFEX – III) reentry vehicle (now

RE-usability Flight Experiment (REFEX)).

SHEFEX-III Experiment: The SHEFEX-III

flight experiment was the third in a series of flight

experiments designed to study the use of sharp

edges and flat plates on hypersonic reentry vehicles

(as opposed to the traditional blunt-body design).

SHEFEX-III was slated for launch in 2016, and

was designed as a spaceplane-like body for lift-

assisted reentry (Fig. 1). The mission profile was

then altered and the mission was renamed to

REFEX and is now scheduled for flight in 2019 [3].

(a)

(b)

Figure 1: The SHEFEX III reentry vehicle. (a) Front

view [2](b) Side view

SHEFEX-III Boundary Layer Transition

Prediction Tool: A boundary layer transition

prediction tool was created to produce estimates of

the transition location on the windward surface of

the SHEFEX-III reentry vehicle, given a flight

profile. Recently, Kolbe et al. [4] published an

analysis of heat loads on the SHEFEX-III vehicle.

That work presented a predominantly inviscid

calculation of heat loads. Further, the evolution of

boundary layer parameters or transition location

was not included in their analysis. The boundary

layer parameters and transition locations predicted

by the methods presented in our work can also be

used to produce analytical estimates of heat loads

on the vehicle.





Methods: The ramp-like shape of the

windward (bottom) surface of the SHEFEX-III

reentry vehicle lends itself to theoretical analysis

based on shock-expansion theory. Fig. 2 shows the

inviscid flow profile over the vehicle. The

interaction of the flow with the first ramp causes a

shock to appear at the nose of the vehicle. This

shock may be an attached oblique shock or a

detached bow shock based on the freestream Mach

number and the angle of attack. The properties

downstream of the shock may be determined using

oblique-shock relations where applicable, and the

properties of the flow downstream of the ramp

junction may be generated using the Prandtl-Meyer

relation for expansion fans [5].

Figure 2: Inviscid flow structure on the bottom surface

of SHEFEX-III

With the inviscid properties calculated, the

boundary layer parameters must now be estimated

for a compressible boundary layer. For this task,

the Meador-Smart Reference Temperature Method

[6] is used to first produce estimates of relevant

boundary layer quantities such as various boundary

layer thicknesses and Reynolds number based on

position. The wall is assumed to be at radiative

equilibrium with its surroundings. A patch

condition is required in order to extrapolate

relevant boundary layer parameters over the

junction on to the second ramp. The von Karman

integral momentum equation [7] is integrated

across the junction of the two ramps for the

incompressible boundary layer parameters and

then transformed to their compressible

counterparts using the Meador-Smart method.

With the relevant boundary layer parameters

calculated, various empirical correlations were

used to compute the boundary layer transition

location along the bottom surface. Correlations

used for this study were – ReΘ-Me correlation,

Simeonides correlation [8], Reda Correlation [9]

and the Reshotko correlation [10]. With the

transition location obtained, integral properties

such as wall heat flux and friction coefficients

could now be calculated. The main advantage of

this analysis lies in the fact that correlation

quantities can be generated in merely 2 seconds per

flight point on a laptop computer, as opposed to a

Navier-Stokes solution which may take hours to

converge. If the correlations are themselves

empirical, the extra computation time for a full

Navier Stokes solution is not necessarily

advantageous. The time saved by using the analysis

method mentioned above can be used to identify

interesting cases for a more detailed analysis.

Tool Validation: The Shefex-III transition

prediction tool was validated using 2-D CFD

analysis of the reentry vehicle at four points in the

flight profile. This analysis was performed using

the DLR Tau Code. The flight points were selected

over a wide range in flow conditions – from low

angle of attack, low Mach number flows to high

Mach number non-equilibrium flows.

Projected Results: The final work presented

at the conference will include a more detailed

description of the process used to estimate

boundary layer quantities. Tool validation will be

discussed in greater detail, including full

specification of the test cases chosen and a

comparison between predictions from the

SHEFEX-III tool and from Navier-Stokes

solutions using the DLR Tau Code. These results

will include predictions of correlation quantities

and integrated coefficients such as the heat flux

coefficient and the coefficient of friction.

References:

[1] S. P. Schneider, "Laminar-Turbulent

Transition on Reentry Capsules and

Planetary Probes," Journal of Spacecraft and

Rockets, vol. 43, no. 6, 2006.

[2] "DLR Press Portal: Mechanical tests for

SHEFEX," German Aerospace Center -

Licensed under Creative Commons License,

[Online]. Available:

http://www.dlr.de/dlr/presse/en/desktopde



fault.aspx/tabid-10172/213_read-

3330/#/gallery/1343.

[3] A. Guelhan, Personal Communication, 2016.

[4] A. Kolbe, C. Dittert, T. Reimer, H. Weihs and

V. Watermann, "Analytical Determination of

the Thermal Loads on the re-entry vehicle

SHEFEX III," in 8th European Symposium on

Aerothermodynamics for Space Vehicles,

2015.

[5] J. D. Anderson, Hypersonics and High-

Temperature Gas Dynamics, AIAA Education

Series, 2006.

[6] W. E. M. a. M. K. Smart, "Reference Enthalpy

Method Developed from Solutions of the

Boundary-Layer Equations," AIAA Journal,

vol. 43, no. 1, 2005.

[7] H. Schlichting, Boundary Layer Theory,

McGraw-Hill Book Company, 1979.

[8] G. A. Simeonides, "Correlation of laminar-

turbulent transition data over flat plates in

supersonic/hypersonic flow including leading

edge bluntness effects," Shock Waves, vol.

12, pp. 497-508, 2003.

[9] D. C. Reda, "Review and Synthesis of

Roughness-Dominated Transition

Correlations for Reentry Applications,"

Journal of Spacecraft and Rockets, vol. 39,

no. 2, 2002.

[1

0]

E. Reshotko, "Transition Prediction:

Supersonic and Hypersonic Flows," NASA-

OTAN Report RTO-EN-AVT-151.



HYPERSONIC ENTRY TRAJECTORY OPTIMIZATION OF HIGH BALLISTIC COEFFICIENT

VEHICLES. Christopher G. Lorenz1 and Zachary R. Putnam2, 1University of Illinois at Urbana-Champaign (clo-

[email protected]), 2University of Illinois at Urbana-Champaign ([email protected])

Future entry, descent, and landing systems at Mars

will be required to land larger masses than current state

of the art systems are capable of accommodating. New

advances including supersonic retropropulsion will be

required to land the next generation of vehicles such as

the Mars Sample Return lander and future manned mis-

sions on the Martian surface.

With the exception of the Mars Science Laboratory

(MSL), all missions launched to Mars have used full

lift up trajectories to increase their terminal descent

initiation altitude. [1] MSL was the first lander to use

bank modulation of the lift vector to both achieve bet-

ter final altitude and landing precision performance.

This work focuses on the optimization of hyperson-

ic entry trajectories leading into the terminal superson-

ic retropropulsion phase of flight. A bang-bang control

method has been implemented to optimize these entry

profiles over wide set of initial parameters. These pa-

rameters include vehicle characteristics such as ballis-

tic coefficient and lift to drag ratio as well as trajectory

characteristics including initial entry velocity, initial

flight path angle, and velocity at terminal descent ini-

tation. This trade space is being examined to enable a

better understanding of the entry guidance system re-

quired for these missions.

These trajectories have been analyzed to identify

the point at which the bank modulation of the lift vec-

tor becomes useful (leads to higher altitudes) across the

full range of parameters.

In addition, these trajectories have been analyzed

for compliance with the g-loading constraints of both

manned and unmanned craft. Early results from this

work demonstrate the difficulties of landing very high

mass systems (ballistic coefficients in the range of 450-

600 kg/m2) over a traditional range of entry parame-

ters. This analysis shows that future work may be re-

quired to identify trajectories that meet the human sys-

tems intergration requirements identified during the

Constellation program. [2]

Future work includes studying the interplay of the

hypersonic portion of these trajectories with the super-

sonic retropropulsion terminal descent in concert with

partners at the Georgia Institute of Technology. The

sensitivies of the findings of this work to uncertainties

in atmospheric density and vehicle parameters will also

be examined.

[1] Robert D. Braun and Robert M. Manning.

"Mars Exploration Entry, Descent, and Landing Chal-

lenges", Journal of Spacecraft and Rockets, Vol. 44,

No. 2 (2007), pp. 310-323. [2] Stephenson, Gary V.

"Human Rating Requirements for Mars Missions."

(2012).



Rocket Deployed Stability Flight Testing for a Candidate Aeroshell Shape. A. L. Strahan1 and R. R. Sostaric2, 1NASA-Johnson Space Center, 2101 NASA Parkway, Houston Tx, 77058, mail code EG4, [email protected] and; 2NASA-Johnson Space Center, 2101 NASA Parkway, Houston Tx, 77058, mail code EG5, [email protected].

Introduction: This paper will document an effort

to investigate and characterize the aerodynamic stabil-ity of a candidate aeroshell shape over a range of Mach, from Super to Sub-sonic, via deployment from a high altitude sounding rocket near apogee, followed by free-fall of the aeroshell test article. This test was per-formed on November 6th, 2015, where the test article was deployed from the SpaceLoft XL Sounding rocket, launched from Spaceport America, and landed at the White Sand Missile Range (WSMR), both in south central New Mexico. The rocket flight was funded by the NASA Flight Opportunities Office, while the de-velopment of the test article was funded by a NASA Game Changing Technologies grant under the Space Technologies and Mission Directorate (STMD) Office. While the sounding rocket flight and deployment were successful, the test article was lost on WSMR, due to tracking and telemetry issues, without recovery of any of the desired data. Still the successful performance of the sounding rocket and its deployment system demon-strates a new and relatively inexpensive capability to acquire super and sub-sonic stability and aerodynamic performance data for a candidate planetary aeroshell shape, in an actual unconstrained flight environment. Unlike static aerodynamic data, it is significantly more challenging to accurately acquire such unconstrained dynamic aero performance in a wind tunnel.

Aeroshell Shape: The proposed shape for this sounding rocket based investigation was defined dur-ing the Maraia development activity at the NASA Johnson Space Center. This activity, running from 2011 through 2015, sought to develop a relatively small, semi-autonomous payload return capsule, that could be delivery to the International Space Station (ISS) by a visiting vehicle, stored inside, then trans-ferred outside through the Japanese Experiment Mod-ule (JEM) airlock and released via a spring loaded de-ployment table mounted on a remote ISS arm. The capsule would then maneuver from ISS, deorbit and land, via parachute, at a remote facility such as the Utah Test & Training Range. The capsule shape and related flight elements were also envisioned as being able to evolve into a system that could support a plane-tary sample return mission, by carrying the sample through the earth entry and landing phase after separa-tion from the vehicle stack, which itself may just be on an earth fly-by trajectory. Additionally this capsule shape could carry a payload through much of the at-mospheric entry at a planetary destination.

The specific shape being investigated was essential-ly that of a Soyuz capsule, which was truncated part way up the ‘gumdrop’ shaped back-shell into a flat upper surface, while retaining the Soyuz outer mold line for the lower back-shell and heat shield. The shape was truncated to maximize volume and still fit out the JEM airlock, while hoping to still enjoy some benefit of the strong mono-stability of the original So-yuz. This shape was demonstrated to be quite stable, oscillating out to attitudes of only +/- 15 to 20 deg. (about one half to one third of that predicted), in a flight test from about Mach 0.8 to 0.5, when the article was flown as a secondary payload and dropped from a high altitude balloon. The balloon was provided by the NASA Wallops Flight Facility and operated by the Columbia Scientific Balloon Facility, flown out of Fort Sumner New Mexico in August 19, 2013 [1].

Test Article for Rocket Flight: It was desired to build on the successful balloon drop by expanding the flight test range to cover up to Mach 3.5 and on down through transonic and subsonic to Mach 0.5 or even lower. To accomplish this larger Mach test range, it was proposed to deploy the candidate capsule shape from a high altitude sounding rocket while near apo-gee, such that the test article would achieve the high supersonic speed simply by free-fall, and then deceler-ate through the desired Mach range until again reach-ing chute deploy conditions. The test article would be the same shape developed by the Maraia program and with the same relative Center of Gravity location, as flown in the balloon test, only scaled down to fit within the custom payload test section of the rocket. This proposal was accepted and funded by the NASA Flight Opportunities Office and the NASA Game Changing Technology development programs in cooperation, and a test article was developed between Feb 2014 and August 2015. The test article included instrumentation (attitude, rate, acceleration, altitude, and ‘heat shield’ temperature); a small cold gas thruster system to aug-ment attitude control if dynamics got too large, or con-versely to provide a small excitation pulse if dynamics were too small; a small flight computer to control thrusters, record data and command events; a recovery system, being a parachute deployed by cold gas; a te-lemetry system and two video cameras to also capture dynamics. The test article was 9.7 inch diameter, 6.4 inch tall and with a mass of about 11 lbs. (figure-1).

Sounding Rocket: The flight test delivery plat-form, unique to this experiment, was a SpaceLoft XL sounding rocket provided by UP Aerospace of Denver



figure-1: Maraia Capsule Test Article Colorado, already on contract as a NASA Flight Op-portunities test platform (figure-2). UP Aerospace had successfully flown 9 previous flights for NASA, carry-ing a variety of NASA and student payloads, however this flight was different in that a payload deployment capability was developed by UP Aerospace and inte-grated into their rocket. The payload delivery system was such to allow the deployment of the maximum payload size possible, being one that just fit inside the inner diameter of the rocket (about 10 inches). This was accomplished by a sequence of events occurring after rocket motor burn out and up in the near vacuum of the upper atmosphere (above 250,000 ft). This se-quence involved first deploying the nose cone for a separate recovery, then deploying the test article out of the top of the rocket, through the large opening left by the absent nose cone.

figure-2: SpaceLoft XL Sounding Rocket

The Flight: The rocket was launched at about 8 am on Nov 6, 2015 reaching an apogee of about 119 km

(385,000 ft), where it successfully deployed the Maraia test article a few seconds after deploying its nose cone (figure-3). The limited available telemetry indicated that the capsule survived ascent, woke up and began to function, though for unknown reasons deployed its parachute prematurely and appears to have drifted off its predicted trajectory. Its telemetry system operated only intermittently and WSMR radar tracking, as well as an on-board beacon, did not function such as to al-low recovery of the test article.

figure-3: Maraia at altitude, inside rocket as viewed

from departing nose cone, prior to its deployment Again, while the capsule was lost, this capability,

developed by UP Aerospace in support of the Maraia flight test, now exists on a NASA sponsored platform to deploy any similar sized payload (up to 10 inch di-ameter, 20+ inches tall and 20 to 30 lbs), that could support other candidate EDL flight test shapes, provid-ing 2+ minutes of near vacuum micro-g flight, fol-lowed by atmospheric flight over a range of Mach up to 3.5+, offering the opportunity to investigation stabil-ity and performance.

Summary: This paper will go into more detail re-garding the specific test objectives of this flight, of how the aeroshell shape was adapted, tested and inte-grated with the rocket, and to survive the high stress ascent environment, and to then perform its target mis-sion. Furthermore this paper will describe the events leading up to the flight test, and finally the test out-come including what is believed to have gone wrong with the test article resulting in its loss. Also the paper will provide some lessons learned that could enable future candidate planetary vehicle technologies (aero-shells or whatever) to successfully take advantage of a sounding rocket test platform to explore high altitude and supersonic through subsonic flight, typically the most dynamically challenging portion of entry.

References: [1] Sostaric, R. R. and Strahan, A. L. (2016)

Maraia Capsule Flight Testing and Results for Entry, Descent and Landing, AIAA SciTech Conference, 4-8 Jan 2016, San Diego, CA; Document ID 20150021959.



Dynamic stability of ERC via balloon drop tests . J.S. Lingard

1, J.C. Underwood

1, S.J. Overend

1, A. Saunders

1, E. Johnstone

2, D. Rebuffat

3

1Vorticity Ltd, Chalgrove, Oxfordshire, OX44 7RW, United Kingdom [email protected]

2Fluid Gravity Engineering, Emsworth, Hampshire, PO10 7DX, [email protected]

3European Space Agency, Keplerlaan 1, 2201 AZ Noordwijk, Netherlands, [email protected]

This paper will discuss a recently completed pro-

gramme performed under ESA MREP-2 to investigate

the aerodynamic properties of three different Earth

Return Capsules (ERC) by balloon drop testing.

One of the long-term goals of space exploration is

the Mars Sample Return (MSR) mission. The final part

of this mission will use an ERC to return the sample to

Earth. There is also interest in an asteroid sample re-

turn mission within the Cosmic Vision programme.

These missions foresee a capsule re-entering the

Earth’s atmosphere without the use of a parachute. In

this scenario, the stability of the ERC during entry and

descent is vital.

There are two types of aerodynamic stability: static

and dynamic. A statically unstable vehicle that receives

a perturbation in angle of attack will experience a pitch

moment that acts to increase the angle of attack.

A statically stable vehicle will experience the oppo-

site. This stabilising moment results in a series of si-

nusoidal oscillations. If the vehicle is dynamically sta-

ble the magnitude of these oscillations will decrease. If

the vehicle is dynamically unstable, the magnitude will

increase. Dynamic stability varies as a function of

Mach number, angle of attack and centre of gravity

position.

Measuring dynamic stability in a wind tunnel is dif-

ficult because the support structure has an impact on

the flow. Free-flight testing using ballistic ranges has

been conducted but this testing occurs at non-

representative Reynolds numbers and Reduced Fre-

quency Parameters (RFP) and flight times are very

short..

Vorticity developed a test technique to investigate

the dynamic stability of three different candidate ERC

geometries, at representative Reynolds numbers and

Reduced Frequency Parameters in the critical transonic

Mach range. A full scale model of each vehicle was

built to cover Mach numbers up to the high subsonic

and a ~25 cm subscale vehicle was built to cover Mach

numbers between 0.7 and 1.5. These vehicles were

flown to 30-35 km under helium balloons over North-

ern Sweden, released and allowed to free-fall to the

ground. The sub-scale vehicle was housed within an

aerodynamic fairing for the initial part of the trajectory

in order to allow a supersonic velocity to be achieved.

On each vehicle, miniaturised instrumentation was

used to record accelerations, rates and position which

were then used to reconstruct the trajectory of the vehi-

cle: a coherent set of positions and orientations; linear

and angular velocities; and linear and angular accelera-

tions. The instrumentation was ejected from the test

vehicles just before impact and descended to the

ground under a small parachute.

The trajectories were then analysed to determine an

aerodynamic database (both static and dynamic) for

each of the vehicles. Axial and normal force coeffi-

cients, the pitch moment and pitch damping coefficient

(Cmq) were determined from the reconstructed trajecto-

ries.

This has led to an improvement in the aerodynamic

database in terms of the subsonic and transonic axial

force coefficients, normal force coefficients and pitch

damping coefficients.

The paper will describe the ERC shapes tested, the

test technique and the resulting aerodynamic databases.




Dynamic CFD Simulations of the Supersonic Inflatable Aerodynamic Decelerator (SIAD-R) Ballistic Range Tests. J. M. Brock and E. C. Stern2, 1Analytical Mechanics Associates, Inc., Moffett Field, CA 94045. Corresponding author: [email protected], 2NASA Ames Research Center, Moffett Field, CA 94045. Corresponding author [email protected]

A method for simulating the unsteady dynamics of entry vehicles during the descent phase of EDL has been implemented within the US3D computational fluid dynamics (CFD) solver. Currently, the dynamic stability of blunt-body entry vehicles must be deter-mined experimentally, which makes it difficult to ex-amine the physical mechanisms which drive dynamic stability. With a high-fidelity CFD tool capable of sim-ulating dynamics of the entry vehicle, further interro-gation into the flow physics is possible at reduced cost. While previous proof-of-concept studies [1] have shown this novel approach to have promise, more rig-orous validation is necessary if this method is to be incorporated into the aerodynamic design process. The primary objective of this work is the initial validation of the solver through direct comparison against exper-imental data of an existing aerodynamic decelerator technology development effort[2].

The presentation will be split into two parts. The first half will discuss the methodology behind perform-ing dynamic simulations of complex flow topologies. Here, the basics of the US3D flow and dynamics solvers, key aspects of grid-generation, and the effect of numerics will be presented and discussed. In partic-ular, the importance of adequately resolving key fea-tures of the flow (i.e. the wake) are emphasized. The first part will conclude with an overview of the analy-sis workflow required to carry out a single dynamic simulation from start to finish.

The second half of the presentation will discuss the comparison of the simulated results to experimental values. A brief overview of the experimental set-up at the Hypervelocity Free-Flight Aerodynamics Facility (HFFAF) located at the NASA Ames Research Center (ARC), will be presented. A selection from the results of the experiment [2] will be used to compare against predicted dynamic behavior from CFD. The purpose of this section will be to highlight similarities and differ-ences in the predicted results of US3D against the experiment. Computed trajectories and static/dynamic aerodynamic coefficients will be presented and their comparisons to experiment discussed.

[1] Stern, E. C., et al., “Estimation of Dynamics Stability Coeffi-cients for Aerodynamic Decelerators Using CFD”, AIAA Paper, New Orleans, 2012. [2] Wilder, M., Bogdanoff, D., Brown, J., Yates, L., “Aerodynamic Coefficients from Aeroballistic Range Testing of Deployed and Stowed SIAD SFDT models”, NASA TM to be published 2016.

Top: Comparison of experimental[2] Schlieren image (left) and pseudo-Schlieren using density gradient magnitude from CFD (right) Middle: Temperature contours to show wake structure from CFD. Bottom: Angle of attack vs time for experiment and CFD.





Preliminary Results of a Ballistic Range Test of a 2% Scale Mars 2020 Entry Capsule with Onboard Pressure Measurement. M. Schoenenberger1 , T. G. Brown2 and L. A. Yates3 1NASA Langley Research Center, Hampton, VA 23681, 2U.S. Army Research Laboratory, Aberdeen MD, 2AerospaceComputing Inc., Mountain View CA

Abstract: Preliminary results of a 2% scale

supersonic ballistic range test of the Mars 2020 entry capsule with onboard pressure instrumentation are presented. The test was led by NASA Langley Research Center and designed and conducted in partnership with the U.S. Army Research Laboratory, at their Transonic Experimental Facility in Aberdeen Maryland. The primary objective of the test was to obtain backshell surface pressure measurements at supersonic free-flight conditions. The data will be used to help select the pressure port location (or locations) where a transducer will be installed on the Mars 2020 entry capsule. Four models were constructed and each were instrumented with four pressure transducers. Pressure transducers measuring stagnation and backshell pressures were used together with shadowgraphs taken in the range to reconstruct how the surface pressures vary with changes in total angle of attach and Mach number.

Note: Ballistic range testing is scheduled to occur around the time this abstract is submitted. No results are included here.

Introduction: The Mars Science Laboratory (MSL) entry capsule successfully flew through Mars’ atmosphere in 2012, delivering the Curiosity Rover safely to the surface. Mounted to the MSL forebody heatshield was the MEADS experiment. Part of the MSL Entry Descent and Landing Instrumentation (MEDLI) experiment, MEADS consisted of seven pressure transducers that measured surface pressures throughout entry with data being recorded until just prior to heatshield jettison while the capsule was under that parachute. The pressure data returned by MEADS provided excellent information to reconstruct dynamic pressure, atmospheric density and angles of attack and sideslip of the vehicle throughout entry1. Much of the data corroborated and validated preflight aerodynamic predictions that were generated primarily with computational fluid dynamics (CFD) codes. One notable discrepancy was observed at low supersonic conditions. The reconstructed axial force coefficient was much greater than preflight predictions. However, the large differences were observed at dynamic pressures lower than the minimum values for which MEADS was designed. Conclusive reconciliation of preflight predictions with low Mach measurements was not possible.

Part of MEDLI2, the MEADS2 experiment will fly on the Mars 2020 mission. Mars 2020 will land

another MSL-class rover using an aeroshell identical to MSL. MEADS 2 is being designed to obtain accurate pressures in the low supersonic regime to resolve the observed discrepancies and reconstruct the axial force coefficient near parachute deploy conditions. To fully reconstruct the axial force at low supersonic conditions, the contribution from pressure on the backshell must be measured. CFD prediction and wind tunnel data indicate that most of the wake flow should be separated and roughly constant over most of the vehicle backshell. Therefore, a single pressure transducer should be sufficient to resolve the wake contribution to the total axial force coefficient. At large angles of attack some flow on the backshell becomes attached. The smaller attached region will see pressures different than the separated region. The project must select a backshell port location that remains in separated flow for the best data understand the total CA. To validate CFD predictions and investigate the variation of pressures on the backshell surface, it was decided to conduct a ballistic range test with onboard pressure instrumentation. The pressure data is expected to reveal any flow phenomena that might require more than a single backshell pressure measurement to meet the MEADS2 accuracy requirements for reconciling the reconstructed Mars 2020 axial force coefficient with predictions. Also, the ballistic range test will measure the vehicle Mach number and velocity independent of the onboard instrumentation. This test will validate and quantify the accuracy of the methodology of using of a single backshell transducer to determine the backshell contribution to CA.

Experiment Design: The ballistic range test will be conducted at the Transonic Experimental Facility (TEF), run by the U.S. Army Research Laboratory, Aberdeen MD. The TEF consists of a 200m indoor hallway, instrumented with 5 groups of 5 orthogonal spark shadowgraphs stations. Infrared light triggers initiate each spark source as the model flies down the range. Models are launched down the range using a 120mm smooth bore powder-charge artillery gun. The orthogonal shadowgraphs record the capsule position and orientation vs. time. Post test reconstruction will fit 6-degree-of-freedom trajectories through the position/orientation data to provide continuous velocity and attitude data for comparison to onboard pressure instrumentation. Figure 1 shows an example



shadowgraph of an MSL capsule flying through the TEF range.

The projectiles to be tested are 90mm diameter scale models of the Mars 2020 capsule. They are fabricated from a combination of steel and tungsten. The material distribution was selected to locate the axial cg at the station of the Mars 2020 flight vehicle. (nominally xcg/D = 0.30). The models are instrumented with a suite of magnetometers, accelerometers, and rate sensors along with the primary instrumentation, Kulite XCEL-072 pressure transducers. Each model has a pressure transducer with a surface port located at the vehicle nose (stagnation point at trim conditions) and three transducers on the vehicle backshell. The forebody transducer has a full-scale range of 0-500psi, while the backshell transducers are 0-25psi.

Four models were built for the test. Figure 2 shows one of the fully assembled models, installed in a moment-of-inertia fixture. Two different transducer arrangements were selected and two of each were built. Figure 3 shows the candidate locations available with the onboard instrumentation package. The port locations that are populated for each configuration are also indicated. One backshell transducer on one of the Configuration 1 models was plugged with epoxy to ensure that no external pressure reached that transducer. This was done to determine if the accelerations of the model experienced during launch and flight downrange show up in the pressure signal. The model with a blocked port is designated configuration 1a, a variant of Configuration 1.

Test Plan: Preliminary outdoor ballistic range shots indicated that the models can swerve by four feet or more, from the aim point at the end of the range, if there are small asymmetries (OML geometry or mass distribution) that result in a no-zero trim angle. Analysis was done to determine the maximum allowable cg offset required to keep swerve small enough so the models can be caught downrange with a stack of catcher devices. The data must be downloaded off the models after the shots, so they must be recovered intact and functional. The preliminary outdoor qualification testing determined the downrange catcher location (~170m from the gun barrel exit), initial Mach number (Mach=2.50), and model catcher area required for successful capture of the models. Quality Assurance (QA) measurements of the vehicle OML and mass properties measurements were done to confirm that the model centers of gravity were within the minimum offset value. Combined, these test design parameters should ensure reasonable success in capturing the models for data download.

The test plan calls for each of the four models to be launched twice for a total of eight shots. Minimum

success will be achieved if one model is recovered and data is downloaded and correlated with the 6-DoF trajectory reconstruction. The Mach number will be measured using the reconstructed velocity and stagnation pressure histories. The backshell surface pressures will be compared to preflight predictions determined with a base pressure model created by Mitcheltree2 using Mars Viking entry vehicle data.

Figure 4 shows the stagnation and backshell pressures vs. Mach. The stagnation pressure was calculated using the Rayleigh-pitot equation. The backshell pressure prediction using the Mitcheltree model2 is also plotted with estimated error bounds. The ballistic range experimental data will be compared to these predictions and any variation with angle of attack will be determined by correlating the pressure history to the 6-DoF trajectories reconstructed through shadowgraph data.

Results: The test is currently planned for the week of March 28, 2016 or the week following. Data will be presented at IPPW16

References: [1] Schoenenberger, M. et al. JSR, 51, pp. 1076-

1093. [2] Dyakonov, A. A. et al. AIAA 2012-2999, 2012 ��.

Figure 1 Example shadowgraph of MSL model in ARL TEF



Figure 2 Instrumented ballistic range model installed in moment-of-inertia fixture.

Figure 3 Mars 2020 instrumented ballistic range model with pressure transducer locations (population options and as installed in configurations 1, 1a and 2)

Figure 4 Predicted stagnation and backshell pressures to be measured across ballistic range test conditions.





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