Bryan Palaszewski

Glenn Research Center, Cleveland, Ohio

Entry, Descent, and Landing with Propulsive 

Deceleration: Supersonic Retropropulsion 

Wind Tunnel Testing and Shock Phenomena

NASA/TM—2014-217891

September 2014

AIAA–2013–0027
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Bryan Palaszewski

Glenn Research Center, Cleveland, Ohio

Entry, Descent, and Landing with Propulsive 

Deceleration: Supersonic Retropropulsion 

Wind Tunnel Testing and Shock Phenomena

NASA/TM—2014-217891

September 2014

AIAA–2013–0027

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

51st Aerospace Sciences Meeting

sponsored by the American Institute of Aeronautics and Astronautics

Grapevine, Texas, January 7–10, 2013
Acknowledgments

We would like to thank Bob McCluskey (NASA GRC), Steve Pesek (Jacobs Sverdrup, JSV), Barry Piendl (JSV), and Scott Smrdel 

(also of JSV). The team also included the test hardware design expertise of Jim Buckley (Arctic Slope Research Corporation, ASRC) 

and David Root (also of ASRC). Also many thanks go out to Tim Bencic (NASA GRC) for work in pressure and temperature 

sensitive paints and the NASA GRC Imaging Technology Team of Chris Lynch (Wyle Information Systems, LLC) and Quentin 

Schwinn (Wyle Information Systems, LLC) for their expertise with high speed video camera setup and video data capturing. 

Available from

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7115 Standard Drive

Hanover, MD 21076–1320

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

Level of Review: This material has been technically reviewed by technical management. 
NASA/TM—2014-217891 1

Entry, Descent, and Landing with Propulsive Deceleration: 

Supersonic Retropropulsion Wind Tunnel Testing

and Shock Phenomena

Bryan Palaszewski

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135

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 moon atmospheres for entry, and 

descent 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 retropropulsion (SRP) rocket system for the final soft landing.

A three engine retropropulsion configuration with a 2.5 in. diameter sphere-cone aeroshell model was 

tested in the NASA Glenn Research Center’s 1- by 1-ft (1×1) 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 SRP system. This 

baseline testing defined the flow field around the aeroshell and from this comparative baseline data, 

retropropulsion options will be assessed. Images and analyses from the SWT testing with 300- and 

500-psia rocket engine chamber pressures are presented here. In addition, special topics of 

electromagnetic interference with retropropulsion induced shock waves and retropropulsion for Earth 

launched booster recovery are also addressed.

1.0 Introduction

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

body in the solar system which possesses an atmosphere. Thusly, 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 and of course the final touchdown. In the outer planet atmospheres, the 

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

2.0 Historical Missions

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

guidance. As there is no atmosphere surrounding Earth’s Moon, the lunar landings of 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 more rugged areas of Mars, an additional airbag system was devised for the Pathfinder and 

MER landers to assure a successful landing in rock strewn sites.
NASA/TM—2014-217891 2

3.0 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 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 and 3), parachutes are 

impractical for vehicles with lander masses of over 20 metric tons. The lander’s parachutes would be too big 

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

supersonic speeds) and SRP has been suggested. Many past studies have investigated landing on Mars with 

aerodynamic systems (Ref. 4 to 8). However, the most recent studies imply that the past studies assumptions 

are too optimistic and are in need of revision to assure success. Supersonic retropropulsion, perhaps 

beginning as early as Mach 5, will therefore likely be required for soft landing on Mars.

4.0 Experimental Planning

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

of many tens of tons will likely require more energetic retropropulsion. Figure 2 and Figure 3 show some 

of the historical testing on SRP (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 SRP information, a series 

of test programs were established and planned. The 1×1 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.

Model development began for a 2.5-in. diameter aeroshell. The 2.5-in. 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 from 0° to 20° of angle of attack (AoA). The construction of 

the model and sting strut was stainless steel. The model was also instrumented with both temperature 

sensors and pressure transducers. There were three thermocouples and nine pressure ports on the windward 

side of aeroshell. There were three thermocouples and three pressure ports on 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.

5.0 Test Data

Testing commenced on March 17, 2010, for a 1 day period. The tunnel operations were very smooth 

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

shock was established. The tunnel 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. Trailer-provided 

air was used for the simulated rocket engine flow. The rocket nozzle design was derived from Reference 9.

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. Based on previous testing, no unanticipated aerodynamic blockage 

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

unstarts occurred in only several runs, and their occurrences are noted in Table I. The tunnel unstarts 

occurred with all 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.
NASA/TM—2014-217891 3

Figure 4 shows a typical Schlieren image for the baseline SWT testing three retropropulsion 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. Additional data was gathered at Mach 

Number = 3.5, 4.0, and 5.0, with the angle of attack at 0.0°, and these results are shown in Figure 4 to 

Figure 7, respectively. As the Mach number increases, there is less noticeable or no wall shock interaction 

in the images. On most runs, we are searching and striving to reach the lowest Reynolds number (Re)/foot 

and the lowest total pressure at each Mach number, so that we can more accurately simulate the Mars 

entry conditions. Higher values of Re/foot can represent other atmospheric entries into Earth, the outer 

planets, and 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. 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 retropropulsion 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):

Where:

CT Thrust coefficient

γ∞ Ratio of specific heats at infinity

M∞ Mach Number at infinity

Pe Pressure at nozzle exit

P∞ Pressure at infinity (tunnel pressure)

Ae Nozzle exit area

AB Test article projected area

γe 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.

Testing with retrorocket configurations was planned to include flexible changes of the nozzle 

expansion ratio and the angle of attack. The overall design of a retropropulsion model is shown in 

Figure 9 to Figure 11. Three expansion rations of 10:1, 20:1 and 4:1 are shown, respectively. Appendix A

shows 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° . Again, 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 Re/foot. Appendix C provides 

a detailed drawing of the windward side of the aeroshell test model.

6.0 Thoughts on Alternate Retropropulsion Configurations

Figure 12 shows the Mars Lander configuration of the Design Reference Architecture 5 (DRA-5, 

Ref. 8). As currently designed, the vehicle has a large series of open trusses that support the subsystems 

of the vehicle: tankage, propellants, engines, rover(s), return vehicle, etc. During EDL, it has been 

suggested that the aeroshell surrounding the vehicle can be released at supersonic speed and the main 

engines be used for SRP. With all of the major open trusses of the lander structure in the aerodynamics 
NASA/TM—2014-217891 4

stream, this would lead to severe damage to the lander (and is not recommended). By using a combination 

of deployable structures and SRP, the vehicle could be much more controllable and safe from unwanted 

aerodynamic heating. 

Saturn I retrorocket separation—The Saturn I rocket from the 1960s used a retrorocket system to 

assist with the first and second stage separation. The separation motors were solid rockets, and were used 

to assure that the stages did not collide during separation. The retro rocket flow field was analyzed (see 

Figure 13 and Figure 14) and predictions made of the influence of the flow field on vehicle 

communications (Ref. 10). Such analyses will likely be important in future Mars exploration missions 

using SRP.

U.S. Air Force (USAF) reusable booster system (RBS) Rocketback maneuver—The USAF RBS has 

been suggested as a potential new launch vehicle (Refs. 11 and 12). The RBS is composed on a reusable rocket 

powered first stage and a rocket powered second stage, as shown in Figure 15. In the suggested design, the 

booster staging separation velocity is so high that the first stage must employ a rocket back maneuver (see 

Figure 15). After staging, the first stage vehicle will turn to fire its main engines into the oncoming airflow, 

and slow the vehicle down so that it may return to the launch site. Figure 16 shows a flowfield calculation for 

4 angles of attack (Ref. 12). Severe heating may be experienced during this maneuver.

Fins—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 (Russian Soyuz launch vehicle, etc.). The 4 grid fins 

are mounted on the sides of the vehicles and provide enhanced stability during the use of the launch 

escape system. U.S. Army and international missile testing (Refs. 13 to 23) has 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.)

7.0 Concluding Remarks

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

deceleration (or SRP). Initial data gathering was successful and this data will be used as the comparative 

baseline for upcoming larger scale retropropulsion 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 retrorocket—shock interactions. More detailed data and image 

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

retrorocket 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 (EDL) 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. 24

to 40). 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.
NASA/TM—2014-217891 5

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

[A 2.5-in.-diameter aeroshell model, three engine configuration. NASA EDL SRP 1×1 SWT test results summary, March 17, 2010]

Reading Mach

no.

AoA,

deg.

Pjet,

psia

Comment Specifics

6 2.5 0 200 Some strong wall 

interaction

Some strong wall interactions interaction

9 2.5 0 300 Tunnel unstart, with 

recovery

Post firing: this firing is a potential unstart, with the tunnel taking a few 

seconds to recover the full Mach number after the engines are turned off.

10 2.5 0 500 Tunnel unstart, with 

recovery

Post firing: this firing is a potential unstart, with the tunnel recovering 

the full Mach number flow after the engines are turned off.

7 2.5 10 200 Some strong wall 

interaction

Some strong wall interaction

8 2.5 10 300 Tunnel unstart, with 

recovery

Post firing: this firing is a potential unstart, with the tunnel recovering 

the full Mach number flow after the engines are turned off.

11 2.5 10 500 Tunnel unstart, with 

recovery

Post firing: this firing is a potential unstart, with the tunnel recovering 

the full Mach number flow after the engines are turned off.

19 3.0 0 200 Spectacular Minor wall interactions

20 3.0 0 300 Spectacular Some strong wall interactions

23 3.0 0 500 Tunnel unstart, with 

recovery

Post firing: this firing is a potential unstart, with the tunnel recovering 

the full Mach number flow after the engines are turned off.

16A, 17 3.0 10 200 Spectacular Minor wall interactions

21 3.0 10 300 Spectacular Minor wall interactions

22 3.0 10 500 Tunnel unstart, with 

recovery

Post firing: this firing is a potential unstart, with the tunnel recovering 

the full Mach number flow after the engines are turned off.

24 3.0 15 500 Tunnel unstart, with 

recovery

Post firing: this firing is a potential unstart, with the tunnel recovering 

the full Mach number flow after the engines are turned off.

30 3.5 0 200 Spectacular No wall interactions

35 3.5 0 300 Spectacular No wall interactions

39 3.5 0 500 Spectacular No wall interactions*

31 3.5 10 200 Spectacular No wall interactions

34 3.5 10 300 Spectacular No wall interactions

38 3.5 10 500 Spectacular No wall interactions*

No rdg. 3.5 15 200 No data point taken No data point taken

36 3.5 15 300 Spectacular No wall interactions

37 3.5 15 500 Spectacular No wall interactions*

43 4.0 0 200 Spectacular No wall interactions

48 4.0 0 300 Spectacular Some wall interactions

49 4.0 0 500 Spectacular Minor wall interactions

44 4.0 10 200 Spectacular Minor wall interactions, far downstream

47 4.0 10 300 Spectacular No wall interactions

50 4.0 10 500 Spectacular Minor wall interactions

45 4.0 15 200 Spectacular Minor wall interactions, far downstream

46 4.0 15 300 Spectacular Minor wall interactions, far downstream

51 4.0 15 500 Spectacular Minor wall interactions, downstream, near backshell

63 5.0 0 200 Spectacular No wall interactions. Required heater usage, T= 300 °F to establish 

stable shock on the model. Heater used in all M = 5.0 tests

64 5.0 0 300 Spectacular No wall interactions

65 5.0 0 500 Spectacular Minor wall interactions

62 5.0 10 200 Spectacular No wall interactions

61 5.0 10 300 Spectacular No wall interactions

60 5.0 10 500 Spectacular No wall interactions

57 5.0 15 200 Spectacular No wall interactions

58 5.0 15 300 Spectacular No wall interactions

59 5.0 15 500 Spectacular No wall interactions
NASA/TM—2014-217891 6

Figure 1.—Comparison of Viking-spacecraft-like (sphere-cone) aeroshells for Mars entry

(Ref. 1).

Single nozzle 60° aeroshell model with blunt flow interaction, M∞ = 2.0, CT = 1.1.

Figure 2.—Historical retropropulsion testing (Ref. 3, 1970).
NASA/TM—2014-217891 7

Figure 3.—Historical retropropulsion testing, three engine configuration

(Ref. 3, 1970).

Figure 4.—Schlieren image from 1x1 SWT testing—three engine model, Mach = 3.0, Re/ft = 1.45x106,

and Ptotal (psi) = 8.67, AoA = 0°, 300 psia engine chamber pressure.
NASA/TM—2014-217891 8

Figure 5.—Schlieren image from 1x1 SWT testing—three engine model, Mach = 3.5, Re/ft = 1.86x106,

and Ptotal (psi) = 15.00, AoA = 0°, 300 psia engine chamber pressure.

Figure 6.—Schlieren image from 1x1 SWT testing—three engine model, Mach = 4.0, Re/ft = 2.58x106,

and Ptotal (psi) = 26.13, AoA = 0°, 300 psia engine chamber pressure.
NASA/TM—2014-217891 9

Figure 7.—Schlieren image from 1x1 SWT testing—three engine model, Mach = 5.0, Re/ft = 5.19x106,

and Ptotal (psi) = 92.39, AoA = 0°, 300 psia engine chamber pressure.

Figure 8.—Thrust coefficient versus Mach number for varying engine chamber pressures. 
NASA/TM—2014-217891 10

Figure 9.—Retropropulsion model for three engine configuration, with nozzle extensions 

(expansion ratio = 10:1).

Figure 10.—Retropropulsion model for three engine configuration, with nozzle extensions

(expansion ratio = 20:1).
NASA/TM—2014-217891 11

Figure 11.—Retropropulsion model for three engine configuration, with no nozzle extensions

(expansion ratio = 4:1).

Figure 12.—Mars Lander Configuration, NASA DRA-5 (Ref. 8).
NASA/TM—2014-217891 12

Figure 13.—Retrorocket plume at ignition.

Figure 14.—Comparison of retrorocket plum size and 

telemetry signal strenght SA-5.
NASA/TM—2014-217891 13

Figure 15.—Reusable Booster System Flight Path with Rocket Back Maneuver (Ref. 11).

Steady-state Cp values on vehicle and off-body mid-span plane for four angles of attack, M = 4.5, Q = psf

Figure 16.—Reusable Booster System Cp Predictions for 4 Angles of Attack (Ref. 12)

NASA/TM—2014-217891 15

Appendix A.—Schlieren Images 

A.1 AoA = 10° (15°, in Some Cases), 300 psia Chamber Pressure

Figure A.1.—Schlieren image from 1x1 SWT testing—three engine model, Mach = 3.0, Re/ft = 1.42x106,

and Ptotal (psi) = 8.49, AoA = 10°, 300 psia engine chamber pressure.

Figure A.2.—Schlieren image from 1x1 SWT testing—three engine model, Mach = 3.5, Re/ft = 1.87x106,

and Ptotal (psi) = 15.03, AoA = 15°, 300 psia engine chamber pressure.
NASA/TM—2014-217891 16

Figure A.3.—Schlieren image from 1x1 SWT testing—three engine model, Mach = 4.0, Re/ft = 2.57x106,

and Ptotal (psi) = 25.94, AoA = 10°, 300 psia engine chamber pressure.

Figure A.4.—Schlieren image from 1x1 SWT testing—three engine model, Mach = 5.0, Re/ft = 5.32x106,

and Ptotal (psi) = 90.37, AoA = 10°, 300 psia engine chamber pressure.
NASA/TM—2014-217891 17

A.2 AoA = 0°, 500 psia Chamber Pressure

Figure A.5.—Schlieren image from 1x1 SWT testing—three engine model, Mach = 3.0, Re/ft = 1.50x106,

and Ptotal (psi) = 8.95, AoA = 0°, 500 psia engine chamber pressure, tunnel unstart.

Figure A.6.—Schlieren image from 1x1 SWT testing—three engine model, Mach = 3.0, Re/ft = 1.45x106,

and Ptotal (psi) = 8.67, AoA = 0°, 300 psia engine chamber pressure.
NASA/TM—2014-217891 18

Figure A.7.—Schlieren image from 1x1 SWT testing—three engine model, Mach = 4.0, Re/ft = 2.60x106,

and Ptotal (psi) = 26.33, AoA = 0°, 300 psia engine chamber pressure.

Figure A.8.—Schlieren image from 1x1 SWT testing—three engine model, Mach = 5.0, Re/ft = 4.91x106,

and Ptotal (psi) = 89.36, AoA = 0°, 500 psia engine chamber pressure.
NASA/TM—2014-217891 19

A.3 AoA = 10°, 500 psia Chamber Pressure

Figure A.9.—Schlieren image from 1x1 SWT testing—three engine model, Mach = 3.0, Re/ft = 1.44x106,

and Ptotal (psi) = 8.66, AoA = 10°, 500 psia engine chamber pressure, tunnel unstart.

Figure A.10.—Schlieren image from 1x1 SWT testing—three engine model, Mach = 3.5, Re/ft = 1.86x106,

and Ptotal (psi) = 15.00, AoA = 10°, 500 psia engine chamber pressure.
NASA/TM—2014-217891 20

Figure A.11.—Schlieren image from 1x1 SWT testing—three engine model, Mach = 4.0, Re/ft = 2.56x106,

and Ptotal (psi) = 26.01, AoA = 10°, 500 psia engine chamber pressure.

Figure A.12.—Schlieren image from 1x1 SWT testing—three engine model, Mach = 5.0, Re/ft = 5.42x106,

and Ptotal (psi) = 90.40, AoA = 10°, 500 psia engine chamber pressure.
NASA/TM—2014-217891 21

Appendix B.—SRP Run Data, 1×1 SWT

NASA EDL SRP 1×1 SWT TEST RESULTS SUMMARY, MARCH 17, 2010

Rdg Scan Total

_sc

Navg Date Time Ptotal,

psia

Pstatic,

psia

Mach

number

Re/ft Pjet,

psia

AOA,

deg.

6 1 1 0 17-Mar-10 10:05:08 5.10347 0.3425 2.41215 1.10×106 200 0

7 1 1 0 17-Mar-10 10:15:37 5.00908 0.335 2.41438 1.08×106 200 10

8 1 1 0 17-Mar-10 10:23:10 5.05017 0.335 2.41961 1.08×106 300 10

9 1 1 0 17-Mar-10 10:34:14 4.98458 0.3305 2.4199 1.07×106 300 0

10 1 1 0 17-Mar-10 10:40:37 5.02854 0.333 2.4207 1.08×106 500 0

11 1 1 0 17-Mar-10 10:47:04 4.98026 0.3285 2.42324 1.07×106 500 10

16 1 1 0 17-Mar-10 11:28:13 8.5857 0.2705 2.90301 1.44×106 200 10

17 1 1 0 17-Mar-10 11:41:31 8.44954 0.2655 2.90478 1.41×106 200 10

18 1 1 0 17-Mar-10 11:47:55 8.44737 0.265 2.90585 1.41×106

19 1 1 0 17-Mar-10 11:51:28 8.47748 0.2655 2.90696 1.41×106 200 0

20 1 1 0 17-Mar-10 11:58:26 8.66597 0.271 2.90794 1.45×106 300 0

21 1 1 0 17-Mar-10 12:07:46 8.49397 0.266 2.907 1.42×106 300 10

22 1 1 0 17-Mar-10 12:15:42 8.66238 0.2695 2.91134 1.44×106 500 10

23 1 1 0 17-Mar-10 12:22:04 8.94823 0.285 2.89584 1.50×106 500 0

30 1 1 0 17-Mar-10 13:17:15 14.9783 0.207 3.46295 1.88×106 200 0

33 1 1 0 17-Mar-10 13:36:17 15.0123 0.205 3.47135 1.87×106 200 10

34 1 1 0 17-Mar-10 13:37:19 15.0393 0.205 3.47261 1.87×106 300 10

35 1 1 0 17-Mar-10 13:43:25 15.0024 0.2025 3.4795 1.86×106 300 0

36 1 1 0 17-Mar-10 13:49:28 15.0258 0.203 3.47887 1.87×106 300 15

37 1 1 0 17-Mar-10 13:55:21 14.8584 0.201 3.47795 1.85×106 500 15

38 1 1 0 17-Mar-10 14:01:20 15.0031 0.203 3.4778 1.86×106 500 10

39 1 1 0 17-Mar-10 14:07:15 15.0038 0.1985 3.49359 1.85×106 500 0

43 1 1 0 17-Mar-10 14:46:17 25.8632 0.1825 3.94839 2.55×106 200 0

44 1 1 0 17-Mar-10 14:52:20 26.8485 0.1915 3.94039 2.66×106 200 10

45 1 1 0 17-Mar-10 15:00:02 26.1978 0.1845 3.94985 2.58×106 200 15

46 1 1 0 17-Mar-10 15:05:52 26.3128 0.1875 3.9411 2.61×106 300 15

47 1 1 0 17-Mar-10 15:11:27 25.9392 0.185 3.94045 2.57×106 300 10

48 1 1 0 17-Mar-10 15:17:25 26.1281 0.185 3.94585 2.58×106 300 0

49 1 1 0 17-Mar-10 15:23:23 26.3289 0.1855 3.94954 2.60×106 500 0

50 1 1 0 17-Mar-10 15:29:02 26.0068 0.182 3.95456 2.56×106 500 10

51 1 1 0 17-Mar-10 15:34:48 26.7106 0.187 3.95427 2.63×106 500 15

57 1 1 0 17-Mar-10 16:48:45 86.9416 0.176 4.94131 5.54×106 200 15

58 1 1 0 17-Mar-10 16:56:23 90.1885 0.1675 5.01495 5.58×106 300 15

59 1 1 0 17-Mar-10 17:02:19 89.7116 0.161 5.04444 5.48×106 500 15

60 1 1 0 17-Mar-10 17:08:35 90.403 0.1535 5.09235 5.42×106 500 10

61 1 1 0 17-Mar-10 17:15:12 90.3647 0.1455 5.13857 5.32×106 300 10

62 1 1 0 17-Mar-10 17:21:26 90.4469 0.14 5.17308 5.25×106 200 10

63 1 1 0 17-Mar-10 17:27:29 90.0932 0.1325 5.21807 5.14×106 200 0

64 1 1 0 17-Mar-10 17:33:12 92.3942 0.13 5.25723 5.19×106 300 0

65 1 1 0 17-Mar-10 17:39:16 89.3562 0.118 5.31368 4.91×106 500 0

NASA/TM—2014-217891 23

Appendix C.—Model Drawing, Front (Windward) of the Aeroshell
NASA/TM—2014-217891 24

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2004.

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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 moon atmospheres for entry, and descent 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 retropropulsion (SRP) rocket

system for the final soft landing. A three engine retropropulsion configuration with a 2.5 in. diameter sphere-cone aeroshell model was tested in the NASA Glenn Research

Center’s 1- by 1-ft 1×1 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 SRP system. This baseline testing defined the flow field around the aeroshell and from this comparative baseline data, retropropulsion options will be assessed. Images and

analyses from the SWT testing with 300- and 500-psia rocket engine chamber pressures are presented here. In addition, special topics of electromagnetic interference with

retropropulsion induced shock waves and retropropulsion for Earth launched booster recovery are also addressed.

15. SUBJECT TERMS

Planetary entry; Propulsion; Rocket engines; Fluid mechanics; Retrorockets; In situ resource utilization; Retropropulsion

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