Bryan Palaszewski

Glenn Research Center, Cleveland, Ohio

Entry, Descent, and Landing With Propulsive

Deceleration: Supersonic Retropropulsion

Wind Tunnel Testing

NASA/TM—2012-217746

December 2012

AIAA–2012–401
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Bryan Palaszewski

Glenn Research Center, Cleveland, Ohio

Entry, Descent, and Landing With Propulsive

Deceleration: Supersonic Retropropulsion

Wind Tunnel Testing

NASA/TM—2012-217746

December 2012

AIAA–2012–401

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

50th Aerospace Sciences Meeting

sponsored by the American Institute of Aeronautics and Astronautics

Nashville, Tennessee, January 9–12, 2012
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Level of Review: This material has been technically reviewed by technical management. 
NASA/TM—2012-217746 1 

Entry, Descent, and Landing With Propulsive Deceleration: 

Supersonic Retropropulsion Wind Tunnel Testing 

Bryan Palaszewski 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

The future exploration of the Solar System will require innovations in transportation and the use of 

entry, descent, and landing (EDL) systems at many planetary landing sites. The cost of space missions has 

always been prohibitive, and using the natural planetary and planet’s moons’ atmosphere for entry, 

descent, and landing can reduce the cost, mass, and complexity of these missions. This paper will describe 

some of the EDL ideas for planetary entry and survey the overall technologies for EDL that may be 

attractive for future Solar System missions. Future EDL systems may include an inflatable decelerator for 

the initial atmospheric entry and an additional supersonic retro-propulsion (SRP) rocket system for the 

final soft landing. As part of those efforts, NASA began to conduct experiments to gather the 

experimental data to make informed decisions on the “best” EDL options. 

A model of a three engine retro-propulsion configuration with a 2.5 in. diameter sphere-cone 

aeroshell model was tested in the NASA Glenn 1- by 1-Foot Supersonic Wind Tunnel (SWT). The testing 

was conducted to identify potential blockage issues in the tunnel, and visualize the rocket flow and shock 

interactions during supersonic and hypersonic entry conditions. Earlier experimental testing of a 70

Viking-like (sphere-cone) aeroshell was conducted as a baseline for testing of a supersonic retropropulsion system. This baseline testing defined the flow field around the aeroshell and from this 

comparative baseline data, retro-propulsion options will be assessed. Images and analyses from the SWT 

testing with 300- and 500-psia rocket engine chamber pressures are presented here. The rocket engine 

flow was simulated with a non-combusting flow of air. 

Introduction 

Entry, descent, and landing are a series of events needed to safely land on the surface of another 

atmosphere-bearing body in the solar system. Mars, Venus, the outer planets, and the outer planet moon, 

Titan, all require technologies that will protect the spacecraft from the high temperatures created during 

the initial hypersonic entry, and finally slow the vehicle from that hypersonic speed into the supersonic 

regime, then to subsonic velocities and to the final touchdown. In the outer planet atmospheres, the final 

landing would be replaced with a buoyancy system such as an airship, balloon, or an aircraft. 

Historical Missions 

Landing space vehicles on other planetary bodies is a challenge in propulsion, precision control, and 

guidance. As there is no appreciable atmosphere surrounding Earth’s Moon, the lunar landings of the 

robotic Surveyor and human Apollo missions used propulsion for the entire descent. The same was true 

for the successful Luna and Lunakhod flights of the U.S.S.R. For Venus with its dense atmosphere, 

landing vehicles used aeroshell and parachute combinations, with crushable elements (balsa wood, etc.) to 

absorb the final landing energy. On Mars, the landing vehicles became more massive and complex 

(Viking, Pathfinder, Mars Exploration Rovers (MER), Mars Science Laboratory (MSL)), and the since 

the atmosphere was very thin, the final landing systems was a combination for aeroshell, parachute and 

retro rockets. To allow landing in the more rugged areas of Mars, an additional airbag system was devised 

for the Pathfinder and MER landers to assure a successful landing at rock strewn sites. 

 
NASA/TM—2012-217746 2 

Mars 

Several EDL configurations are under assessment for Mars. Figure 1 presents the historical 

comparison of the U.S. Mars entry capsules (Ref. 1). The typical 70 cone angle for these configurations 

was selected for high stability and high drag. As the planet’s atmosphere is quite thin, the blunt body can 

provide the needed drag for relatively small payloads of up to 1 metric ton. As the mass of the lander 

vehicle increases, a different set of EDL technologies are required. Based on past studies (Refs. 2 to 4), 

parachutes are impractical for vehicles with lander masses of over 20 metric tons. The parachutes are too 

big to deploy effectively and reliably. Therefore a combination of inflatable decelerators (for hypersonic 

and supersonic speeds) and supersonic retro-propulsion has been suggested. Many past studies have 

investigated landing on Mars with aerodynamic systems (Refs. 5 to 8). However, the most recent studies 

imply that the past studies assumptions were too optimistic and are in need of revision to assure success. 

Supersonic retro propulsion, perhaps beginning as early as Mach 5, will therefore likely be required for 

soft landing on Mars. 

Experimental Planning 

While the Viking-like aeroshell design has proven successful for missions, higher mass missions of 

many tens of tons will likely require more energetic retro-propulsion. Figures 2 and 3 show some of the 

historical testing on supersonic retro-propulsion (Ref. 3). This testing was only pursued with relatively 

small models and did not result in flight test hardware. To expand the relatively small data base of 

supersonic retro-propulsion information, a series of test programs were established and planned. The 

NASA Glenn Research Center’s 1- by 1 Foot Supersonic Wind Tunnel (SWT) was used for the testing. It 

has a wide range of test velocities from Mach 2.0 to 6.0. Several types of data were gathered during the 

testing: surface pressure measurements, surface temperature measurements, and low speed and high speed 

digital Schlieren video movie imaging. 

Models were developed for a 2.5-in. diameter aeroshell. The 2.5-in. diameter size was selected based on 

the previous wind tunnel testing of the aerodynamic blockage of the tunnel. The initial model was based on 

the 70 sphere-cone shape of the Viking entry capsule. It was attached to a sting-strut that was adjustable 

and can hold the model at a flexible angle of attack (AoA) of 0 to 20. The model and sting strut were 

made of stainless steel. The model was also instrumented with both temperature sensors and pressure 

transducers. There were 3 thermocouples and 9 pressure ports on the windward side of aeroshell. There 

were also three thermocouples and three pressure ports on the leeward side of aeroshell. One additional 

thermocouple was placed near the trailing edge of the strut. High frequency pressure transducers (Kulites) 

were used to measure the engine chamber pressures and tunnel wall pressures in three locations. Optical 

access to the test section allowed imaging with low speed and high speed Schlieren video movie recording. 

The high speed Schlieren recordings were made at 500 frames per second. 

Test Data 

In each test run, the tunnel pressure increased until the flow was started on the model and a stable 

bow shock was established. The pressure was then adjusted until the minimum pressure for tunnel 

operation was reached. Data was taken at this point, and then successive data points were taken at the 

remaining Mach Numbers. Measurements were taken at Mach = 2.5, 3.0, 3.5, 4.0, and 5.0. Trailerprovided air was used for the simulated rocket engine flow. The rocket nozzle design was derived from 

Reference 4. Testing commenced on March 17, 2010 for a 1 day period. 

During the testing, it was noted that with the 2.5-in. model, an initial stable bow shock could be 

established at all Mach Numbers. As expected from previous testing, no unanticipated aerodynamic 

blockage occurred when the engines were not firing. When the rocket engines were firing, tunnel unstarts 

occurred on several runs, and their occurrences are noted in Table I. The tunnel unstarts occurred with all 
NASA/TM—2012-217746 3 

of the 500 psia runs at M = 2.5 and 3.0 and with all of the 300 psia runs at M = 2.5. At all other 

conditions, excellent model performance was demonstrated with minimal wall interactions. 

Figure 3 shows a typical Schlieren image for the baseline SWT testing 3 retro-propulsion engines. 

The Mach Number was 2.91 (M = 3.0 range). The angle of attack was 0. Note that at M = 3.0, the bow 

shock has a small interaction with the tunnel walls in the image. Data was gathered in the 11 SWT at 

Mach Number = 3.5, 4.0, and 5.0, with the angle of attack at 0.0, and these results are shown in 

Figures 4, 5, 6, and 7, respectively. As the Mach number increases, there is less noticeable or no wall 

shock interaction in the images. On most runs, we were attempting to reach the lowest Reynolds 

Number/foot and the lowest total pressure at each Mach Number, to more accurately simulate the Mars 

entry conditions. Higher values of Reynolds Number/foot can represent other atmospheric entries: for 

Earth, the outer planets, or Titan. 

The location of the bow shock very close to the sphere-cone model was unforeseen. The rocket 

engines in past testing have used higher engine pressures of up to 1500 psia, and thus the bow shock is 

often far from the body, perhaps one to several entry vehicle diameters away. The lower pressures used 

here were seen to penetrate the bow shock and that shock remained very near the entry body model. Such 

shock locations will have likely significant influence on vehicle heating due to shock impingement, etc. 

An important parameter for the retro-propulsion testing is the thrust coefficient. It is the ratio of the 

thrust of the vehicle to the drag of the vehicle and is computed with this equation (Ref. 3): 

  2 e e

B

e e T 2 1 M

A

A

P

P

M

2 C  





  

 (1) 

where 

CT Thrust coefficient 

Gamma, infinity Ratio of specific heats at infinity 

M, infinity Mach Number at infinity 

Pe Pressure at nozzle exit 

P, infinity Pressure at infinity (tunnel pressure) 

Ae Nozzle exit area 

AB Test article projected area 

Gamma, exit Ratio of specific heats at nozzle exit 

Me Mach Number at nozzle exit 

Figure 8 illustrates the thrust coefficient versus Mach Number for four engine chamber pressures: 

200, 300, 500, and 1500 psia. The engine expansion ratio is 10:1. For the test cases below 500 psia, the 

thrust coefficient is a maximum of 0.36, Only when the chamber pressure is near 1500 psia and near 

M = 2.0 will the thrust coefficient be equal to or greater than 1.0. Computations of the thrust coefficients 

at other planned expansion ratios (4:1, 20:1, and 50:1) show very similar results. 

The retro- propulsion model configurations were planned to easily change the nozzle expansion ratio 

and the model’s angle of attack. The overall design of a retro-propulsion model is shown in Figures 9, 10, 

and 11. Three expansion rations of 10:1, 20:1, and 4:1 are shown, respectively. Appendices A1, A2, and 

A3 show the Schlieren images from the runs with a chamber pressure of 300 and 500 psia, at an angle of 

attack of 0.0, 10.0, and 15.0. Over the entire test program, rocket engine chamber pressures of 200, 

300, and 500 psia were tested with the 10:1 rocket engine expansion ratio. Appendix B provides the test 

conditions for each run: tunnel total and static pressures and the tunnel Reynolds Number/foot. 

Appendix C provides a detailed drawing of the windward side of the aeroshell test model. 
NASA/TM—2012-217746 4 

Thoughts on Alternate Aerodynamic and Fin Configurations 

Due to the severity and large variations of the flow field from the retro rockets, extensions from the 

entry body may be an important option for stability enhancements. Past testing at supersonic speed of fin 

extensions (grid fins, etc.) shows that such configurations can provide the stability enhancements for 

missiles and human rated vehicles (Soyuz, etc.). Figure 12 shows the configuration for the Soyuz launch 

escape system (Ref. 9). The 4 grid fins are mounted on the sides of the vehicles and provide enhanced 

stability during the use of the launch escape system. Additionally, wind tunnel testing for the Orion Crew 

Exploration Vehicle (CEV) was conducted (Ref. 9) with several configurations of fins and testing was 

conducted at up to Mach 2.5. United States (U.S.) Army and international missile testing (Refs. 10 to 21) 

also evaluated grid fins. The missile testing was for long slender missiles, and hence the application may 

be for a more restricted set of higher lift to drag (L/D) EDL configurations (biconic aeroshells, etc.). 

Other configurations using a petal brake (Ref. 22) may also be compatible with the lower lift to drag 

(L/D) SRP configurations. 

Concluding Remarks 

Experimental programs were planned and executed to gather data of supersonic propulsive deceleration 

(or supersonic retro-propulsion). Initial data gathering was successful and this data will be used as the 

comparative baseline for upcoming larger scale retro-propulsion testing. Schlieren imaging was captured to 

assess the successful formation of the bow shock surrounding the aeroshell. In some cases, the shock 

interactions with the SWT walls occurred and were also visualized. The high speed camera video at 500 

frames per second identified the chaotic nature of the retro-rocket—shock interactions. More detailed data 

and image analyses are continuing. Test planning and model development has been conducted for additional 

retro-rocket equipped aeroshells with different area ratio rocket nozzles: 4:1, 20:1, and 50:1. Due to test time 

limitations, the 4:1, 20:1, and 50:1 expansion ratios were not tested. 

Entry, descent, and landing technologies are under development for the high mass Mars Entry system 

(HMMES). Many investigations of aerodynamic deceleration for the outer planets have been conducted 

as well. The challenges for EDL are numerous, especially for inflatable decelerator and the interactions 

that will occur with propulsive deceleration retro propulsion. The high velocities involved in entry and 

descent will require high temperature materials that are flexible for folding into a small volume, but 

reliable when they are deployed to their full diameter. 

Many exciting possibilities are foreseen for Mars and outer planet exploration and exploitation 

(Refs. 23 to 39). The resources of the outer planets may allow fueling of nuclear fusion vehicles and other 

power plants that may be the engine for all of Earth’s energy. Wresting fuels such as hydrogen and helium 3 

from the gas giant planets may be a critical element of outer planet exploration and also flight to the nearby 

stars. The EDL systems will be an integral part of all of these exploration and exploitation scenarios. 

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NASA/TM—2012-217746 14 

TABLE I.—OVERALL RESULTS OF AND COMMENTS ON EDL SRP TEST MATRIX 

[2.5 in. diameter aeroshell model, three engine configuration.] 
NASA/TM—2012-217746 15 
NASA/TM—2012-217746 16 

 
NASA/TM—2012-217746 17 

Appendix A—AoA = 10 (15, in some cases), 300 psia Chamber Pressure 

A1. AoA = 0 500 psia chamber pressure 
NASA/TM—2012-217746 18 
NASA/TM—2012-217746 19 

A2. AoA = 10 500 psia chamber pressure 
NASA/TM—2012-217746 20 
NASA/TM—2012-217746 21 

A3. A0A = 10°, 500 psia chamber pressure 
NASA/TM—2012-217746 22 

 
NASA/TM—2012-217746 23 

Appendix B—SRP Run Data, 11 SWT 

TABEL II.—NASA EDL SRP 11 SWT TEST RESULTS SUMMARY 

 

NASA/TM—2012-217746 25 

Appendix C—Model Drawing, Front (Windward) of the Aeroshell 
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PLEASE DO NOT RETURN YOUR FORM TO THE ABOVE ADDRESS. 

1. REPORT DATE (DD-MM-YYYY)

01-12-2012 

2. REPORT TYPE

Technical Memorandum 

3. DATES COVERED (From - To)

4. TITLE AND SUBTITLE

Entry, Descent, and Landing With Propulsive Deceleration: Supersonic Retropropulsion 

Wind Tunnel Testing 

5a. CONTRACT NUMBER

5b. GRANT NUMBER

5c. PROGRAM ELEMENT NUMBER

6. AUTHOR(S)

Palaszewski, Bryan 

5d. PROJECT NUMBER

5e. TASK NUMBER

5f. WORK UNIT NUMBER

WBS 017533.02.02.04 

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration 

John H. Glenn Research Center at Lewis Field 

Cleveland, Ohio 44135-3191 

8. PERFORMING ORGANIZATION

 REPORT NUMBER

E-18481 

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration 

Washington, DC 20546-0001 

10. SPONSORING/MONITOR'S

 ACRONYM(S)

NASA 

11. SPONSORING/MONITORING

 REPORT NUMBER

NASA/TM-2012-217746 

12. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified-Unlimited 

Subject Categories: 01, 20, 34, and 91 

Available electronically at http://www.sti.nasa.gov 

This publication is available from the NASA Center for AeroSpace Information, 443-757-5802

13. SUPPLEMENTARY NOTES

14. ABSTRACT

The future exploration of the Solar System will require innovations in transportation and the use of entry, descent, and landing (EDL) 

systems at many planetary landing sites. The cost of space missions has always been prohibitive, and using the natural planetary and 

planet’s moons’ atmosphere for entry, descent, and landing can reduce the cost, mass, and complexity of these missions. This paper will 

describe some of the EDL ideas for planetary entry and survey the overall technologies for EDL that may be attractive for future Solar 

System missions. Future EDL systems may include an inflatable decelerator for the initial atmospheric entry and an additional supersonic 

retro-propulsion (SRP) rocket system for the final soft landing. As part of those efforts, NASA began to conduct experiments to gather the 

experimental data to make informed decisions on the “best” EDL options. A model of a three engine retro-propulsion configuration with a 

2.5 in. diameter sphere-cone aeroshell model was tested in the NASA Glenn 1- by 1-Foot Supersonic Wind Tunnel (SWT). The testing was 

conducted to identify potential blockage issues in the tunnel, and visualize the rocket flow and shock interactions during supersonic and 

hypersonic entry conditions. Earlier experimental testing of a 70° Viking-like (sphere-cone) aeroshell was conducted as a baseline for 

testing of a supersonic retro-propulsion system. This baseline testing defined the flow field around the aeroshell and from this comparative 

baseline data, retro-propulsion options will be assessed. Images and analyses from the SWT testing with 300- and 500-psia rocket engine 

chamber pressures are presented here. The rocket engine flow was simulated with a non-combusting flow of air. 

15. SUBJECT TERMS

Planetary entry; Propulsion; Rocket engines; Fluid mechanics; Retrorocket; in situ resource utilization 

16. SECURITY CLASSIFICATION OF: 17. LIMITATION OF

 ABSTRACT

UU 

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32 

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STI Help Desk (email:help@sti.nasa.gov) 

a. REPORT

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443-757-5802 

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