978-1-4577-0557-1/12/$26.00 2012 IEEE

1

Performance of a Conical Ribbon Drogue Parachute in the Wake of a Subscale Orion Command Module

Anita SenguptaJet Propulsion Laboratory

California Institute of Technology 4800 Oak Grove Dr. Pasadena, CA 91109

Anita.Sengupta@jpl.nasa.gov 818-354-3430

Ricardo Machin, Gary Bourland NASA Johnson Space Center

Houston, TX

Ellen Longmire, Mitch Ryan, Erik Haugen

University of Minnesota Minneapolis, MN

Edward White Texas A&M University

College Station, TX

James Ross NASA Ames Research Center

Moffett Field, CA

Jose Laguna, Robert SinclairAirborne Systems Inc.

Santa Ana, CA

Elsa Hennings US Navy

China Lake, CA

Daniel Bissell TSI Inc.

Shoreview, MN

AbstractTen percent of full-scale Orion conical ribbon drogue parachutes in the wake of a 10% command module were tested in a subsonic 3x2.1 m (10x7) cross section, subsonic atmospheric wind tunnel. The motivation behind the test program is to provide a cost-efficient means to provide both parachute performance data in the wake of the Command Module (CM), and a comprehensive computational fluid dynamics validation dataset, including velocity field measurement in the wake. The subscale parachutes are constructed with the full-scale geometric porosity, number of gores and ribbons, using a novel laser cutting technique. The wind tunnel allows full exploration of the parameter space, including CM pitch plane angle, dynamic pressure (Reynolds number), and trailing distance.

TABLE OF CONTENTS1. INTRODUCTION ................................................. 1!2. EXPERIMENTAL CONFIGURATION ................... 2!3. RESULTS ............................................................ 5!4. FUTURE WORK ................................................. 9!5. CONCLUSIONS ................................................. 10!7. ACKNOWLEDGEMENTS ................................... 10!REFERENCES ....................................................... 10!

1. INTRODUCTIONThe Multi-Purpose-Crew-Vehicle (MPCV) Parachute Assembly System (CPAS) is under development for the terminal descent of the Orion Crew Module (CM). The CM and CPAS are derived from the Apollo architecture. The CM is the same shape as Apollo, but larger with a capacity to hold more crew and a suspended mass of over 8700 kg.

The parachute system consists of two mortar deployed drogues and a cluster of three pilot deployed Ringsail main parachutes. One drogue and one main are for system redundancy, to prevent a single point of failure in the terminal descent system. The drogues are 7.01 m (23 ft) conical ribbon canopies (Figure 2). Conical ribbons are known for their robustness (strength) and stability at Mach numbers approaching transonic speeds. Like Apollo, the purpose of the CPAS drogues is to provide CM stabilization, deceleration prior to the main parachute deploy [1][2]. Like Apollo, Gemini, and Mercury before, the Orion MPCV will also land in the water. The maximum expected deployment space of the CPAS ranges from 1436 to 7996 Pa (20 to 167 psf) and Mach 0.15 to 0.72.

The prohibitive cost of full-scale testing, as well as the need for validation data for computational tools, motivates the use of sub-scale testing for CPAS. Structural testing of the full-scale parachutes typically requires the use of streamlined (lawn dart shaped) fore bodies due to limitation of space and suspended mass capability of standard aircraft. As a result the CPAS drogue is never tested over its full deployment space in a representation aerodynamic wake.

Given this shortcoming, it was determined that a subscale test program would be uniquely suited to probing the aerodynamic and structural response of the drogue in a representative wake [3]. Prior experience with subscale parachute testing for extraterrestrial applications: Huygens, Mars Exploration Rovers (MER), and Mars Science Laboratory (MSL) suggest this approach is valid and can be applied to the Earth Entry systems, as well [4][5][6][7]. The goals of subscale program are:

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2! Measurement of drag in the wake of the CM for reefed and unreefed configurations

! Measurement of parachute dynamic behavior ! Quantification of the effect of CM pitch plane

angle ! Quantification of the effect of trailing distance

The resultant dataset can serve to validate Computational Fluid Dynamics (CFD) and Fluid Structure interaction (FSI) simulations of the Orion, as well as providing much needed data on the drag deficit that result from bluff-body wake. The test program establishes a new aerodynamic database for the Apollo-style drogue as a function of trailing distance, reefing ratio, pitch plane angle, and Reynolds number. This will also represent the first measurement of internal pressure distribution for the CPAS conical ribbon drogue, leading to an improved understanding of the load distribution within the textile elements of the parachute.

Results to be presented include: (1) CM aerodynamic forces and moments, (2) parachute-CM aerodynamic forces and moments, (3) parachute drag, and (4) axial wake velocity profiles downstream of the Orion CM. All of the above were explored experimentally as a function of dynamic pressure, CM pitch plane angle, and parachute geometric properties.

Figure 1. Full-scale CPAS drogue from a low-altitude flight test. The parachute is flown in the wake of a stream-lined payload and is shown in the full-open or disreefed configuration[1].

Figure 2. Full-scale drogue planform area [1].

2. EXPERIMENTAL CONFIGURATIONWind Tunnel Configuration The test was conducted in the Texas A&M University Oran W. Nicks subsonic wind tunnel. The tunnel is closed loop, atmospheric pressure, capable of up to 4778 Pa (100 psf) operation. The test section dimensions are 3x2.1 by 5.5m long (10x7x20ft). The tunnel walls and ceiling are clear poly-carbonate providing ample optical axis. The wind tunnel model was mounted from the side wall on a steel strut. The complete test matrix is shown in Table 1.

A 0.49-m maximum diameter CM was constructed with a 30-degree single-angle back-shell geometry and 70 degree half angle heat shield (Figure 1 and 2). The CM was constructed in three pieces: the heat shield (HS) was fabricated from solid aluminum and the back shell (BS) in two pieces from polycarbonate. The HS and BS were joined with tape to provide a smooth and seamless aerodynamic transition at the shoulder. The HS is mounted to a steel strut that is connected to the wind tunnel wall via a rigid mount on the other end. At the CM end of the strut is a six-component force balance. The HS has a solid hollow fitting that slides over the force balance. The balance provides measurement of only the CM loads. A three axis pancake-type load cell was mounted to the rear of the strut on a solid Al fitting. An eyebolt provided connection to the parachute riser. Use of these two force measurements enabled decoupling of parachute drag from CM aerodynamic forces. The strut had a cylindrical cross section strut mounted to the tunnel via a wider mount wall, providing more rigidity. CM angle of attack variation was obtained by clocking the strut in this mount, rotating around the cross stream axis as shown in Figure 3. The axial distance between the CM and parachute skirt is referred to as the trailing distance (x).Trailing distance is normally non-dimensionalized by the capsule maximum diameter d. Increasing x/d is intended to reduce the effect of the wake deficit on the parachute. Variation of x/d was achieved by increasing the length of the parachute riser and moving the CM upstream in the tunnel, leaving the parachute at the same axial position.

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3Figure 3. Schematics of the test configuration.

Figure 4. Photograph of the CM in the TAMU wind tunnel looking from upstream side at the heat shield.

This approach was selected to keep the Particle Image Velocimetry (PIV) and high speed camera placements and calibration unchanged. An airfoil-shaped fairing was placed over the strut connection to the tunnel wall to create a more streamlined body.

Subscale Parachute

Figure 5. (Top) Subscale parachute construction and single gore geometry. (Bottom) Inflated shape from computer model.

Ten percent scale models of the CPAS variable porosity conical ribbon (VPCR) drogue parachute were fabricated with the geometry scaled from the flight article. Subscale parameters are compared to the full-scale article in Table 2. Each canopy has 24 gores and 19.2% geometry porosity. To reduce the complexity of manufacture, a laser cutting approach was used to create the geometric porosity, as an alternative to individual ribbons. The laser cut gore and full constructed shape are shown in Figure 5. The canopy is made from Nylon broadcloth, Spectra suspension lines and verticals, polyester radials, and a Kevlar single riser. Reefing was accomplished by running a spectra cord through the skirt at each suspension line connection, to the scaled reefing line length ratio, for 1st and 2nd stage respectively. Reefing the parachute is effectively reducing its projected area, as a means to limit the inflation load. Reefing reduces to the effective drag area and can also change the stability characteristics of the parachute. The reefing ratio (RR) and area (S) are defined below: [8]

Table 1. Subscale test matrix.

Configuration d (m) Do(m) x/d

""""####(deg)

Qmax(Pa) Remax

Diagnostic

Capsule Only

0.49

N/A 4, 6, 8, 9.5 0, 30, 50

4778

3x106 PIV, IR, balance Unreefed

0.7

6, 8, 9.5 0, 30, 50 3x106 Balance, load cell 1 stage reefed 6, 8, 9.5 0, 30, 50 3x106 Balance, load cell 2nd stage reefed 6, 8, 9.5 0, 30 3x106 Balance, load cell Constrained 6, 8 0 2x106 PIV, Balance, load cell

Parachutelines

Balance

Load Cell

Flow Direction

CM

Airfoil fairing

Strut

Flow Direction

x/d variation

Strut Mount to Tunnel Wall

42

,

,

4 oooodrrd DS

SCSC

RR $%%

Where Cd,o, Do, and So are the unreefed drag coefficient, nominal diameter, and nominal area. Cd,r and Sr are for the reefed parachute. It should be noted that Do is the constructed, not projected diameter, and is the industry standard metric of measuring a parachute and computing its drag performance.

Reinforcements were used in the vent and skirt to increase robustness for multiple uses. Each canopy was flown multiple times and repairs were made as needed. The interface to the CM was a textile (Kevlar) riser and low-friction swivel. The swivel connected to the eyebolt, thereby making the connection to the load cell.

Table 2. Comparison of full-scale and subscale drogue parachute properties [1],[3].

Particle Image Velocimetry Stereo Particle Image Velocimetry (PIV) was used to obtain planar maps of three component velocity vectors downstream of the CM at several x/d locations, with and without the parachute present. PIV is a valuable diagnostic for obtaining computational fluid mechanics (CFD) validation datasets as it is a non-intrusive flow measurement technique with high spatial resolution. The air in the recirculating tunnel was seeded with mineral oil particles (diameter ~ 1 &m) and the beams from a pair of Nd:YAG lasers were formed into overlapping sheets aimed downward through a gap in the tunnel ceiling. A pair of dual-frame cameras with 12-bit dynamic scaled charged coupled diode (CCD) pixel arrays of 2048 x 2048 captured images of light scattered by the seed particles at two time steps. Camera exposures were synchronized with each laser pulse to 0.25ns resolution using a Synchronizer.

All measurement planes (coincident with the streamwise and vertical directions) were located at the midspan of the tunnel test section. Both cameras were mounted on a horizontal rail on one side of the test section and aimed at the measurement plane from upstream and downstream locations. Fields of view were approximately 0.46d x 0.46d, and for most of the runs, they were located slightly above (0.1d) the axis of the CEV model. A calibration image was collected after each test run for which a square dual-plane plate gridded with fiducial marks was placed at the laser sheet location. Possible errors in the calibration due to plate misalignment were corrected based on a routine in the Insight4G software. Tests showed that the amount of correction was typically very small so that the correction itself caused only minimal changes in resulting velocity statistics. Using the Insight4G (64-bit) software, a total of 1400 image sets (4 images/set) were acquired directly to onboard computer memory (RAM) for each test, collected at an image set capture rate of 7.25 Hz. The image sets were batch processed using the same software and post-processed using a custom MATLAB routine. Each velocity vector in the resulting fields represented a square area with dimension 7.7 mm or 0.016d. Mean and RMS velocity statistics were obtained for each test case.

Measurements were obtained at several locations downstream of the 10% scale CM without and with a parachute. The mounting strut, described previously, is in the plane normal to the image. This orientation was selected to minimize the influence of the strut wake on the PIV data plane. Data was collected at locations ranging from x/d=3.5to 7.1, as detailed in Figure 6. One measurement was also made downstream of the parachute vent. The emphasis was on measuring the same location with respect to the parachute mouth, so as to provide a measurement of the inflow conditions into the parachute for future Computational Fluid Dynamics (CFD) validation purposes.

x/d=3.5 x/d=5.5 x/d=7.1

0.5d0.5d

0.5dCapsule Only

Parachute at x/d=6

Parachute at x/d=8Parachute with Capsule

x/d=5.5 x/d=7.1

Figure 6. (Top) Location and size of PIV data plane is indicated by the blue box. Data planes were just upstream of the parachute skirt. (Bottom) Location of PIV data plane for runs when only the capsule was present. Flow direction is from left to right.

Parameter Symbol FullScale Subscale

Parachute Type --- VPCR VPCR Nominal Diameter (m) Do 7 0.7 Number of Gores --- 24 24 Number of Ribbons --- 52 52 Geometric Porosity --- 19.2% 19% Trailing Distance x/d 6 6-9.5 Suspension Line Length Ls 1.5Do 1.5DoReefing Ratio (%) RR 50-70 50-70 Reynolds Number (x106) Re 1-7

5Pressure Measurement

10 static pressure taps evenly spaced on the heat shield were used to measure the pressure distribution (Figure 7). 10 piezoresistive pressure transducers are planned for measurement of the pressure distribution in the canopy interior [9]. Each unit has a footprint of less than 1.6 x 4.7 mm, 0.2 grams, and provides a 0 to 5 psi dynamic range with a temporal resolution of >1kHz. The amplifier is located downstream of the transducer, minimizing point mass distribution on the canopy. Sensor placement is described later.

Parachute Dynamics

Three high-speed and three high-definition (HD) video cameras were used to measure the dynamic motion of the parachute. Two high-speed and two HD cameras were located upstream of the parachute looking into the interior from above and the side. A single high-speed and HD camera were positioned to obtain a profile view of the inflated shape. Reflective targets were applied with thermal adhesive on the canopy interior, exterior, vent, and leading edge, enabling future photogrammetric post-processing of area, shape, trim angle, and angular rates (not discussed in this paper).

Aerodynamic Load Measurement

A 4500-N inline load cell measured parachute drag at 5 kHz, capturing load transients and frequency content. The load cell only measured parachute loads. A 6-axis 1000-N wind tunnel balance block, mounted inside of the CM, was used to measure forces and moments on the capsule at 5 kHz. All data was recorded and time stamped by the wind tunnel data acquisition system. They are also shown in Figure 7.

Figure 7. Photograph of the heat shield interior indicating pressure tap locations, force balance, and load cell attachment.

3. RESULTSCapsule only runs A study was conducted to determine the optimum configuration and method to trip the boundary layer. Tripping the boundary layer ensures the transition to turbulence occurs at the same location on the heat shield, resulting in a more consistent aerodynamic environment. In addition, the Orion heat shield will exhibit a surface

Table 3. Trip dot study test matrix.

Run Design""""####() Re-''''Run""""####()

Trip height

(in) Re-k

1 30 1000 0 0.0099 550 2 30 1000 30 0.0099 550 3 30 1000 50 0.0099 550

4 30 1000 30 0.0099 &0.0055 550

&~250 5 30 300 30 0.0086 400 6 clean 30 n/a 7 clean 0 n/a

8 0 300 0 0.0080 & 0.0099 450

& 600

9 n/a n/a 0 60 grit Distributed grit

10 0 1000 30 0.0099 & .0080 1000

& 800

roughness post peak heating, a result of the PICA ablation process. Therefore, selecting the appropriate trip height can be used to simulate this roughness, ensuring a more flight like representation. A series of computation fluid dynamics simulations, at the model scale, was performed to select the location and height of the dots resulting in the test matrix shown in Table 3. As the CM can fly at a range of angles of attack (") relative to the free-stream, it was necessary to explore this parameter, as well. The goal of the study was to select a single trip dot configuration that could be used for all angles of attack to simplify the test sequence. #

Two terms are used to describe the trip potential. Re' is the Reynolds numbers based on computed momentum thickness (') and velocity at edge of boundary layer (Ue). Values for transition for the CPAS flight regime are expected to be 300 - 1000. Rek is based on average height of roughness elements (k) and edge velocity and a value of 250 to 500 is expected to cause transition. Both relations use the kinematic viscosity (() of the working fluid. [10][11][12]

vkU

vU

ek

e

%

%

Re

Re ''

Load cell

Forcebalance

Pressure Taps

6Self-adhesive trip dots were applied to the heat shield surface for each configuration in Table 3. The "=30, Re'=1000 pattern is shown in Figure 8. The dots were replaced for each run to change height and location. After the trip dot pattern was applied, the heat shield was cooled to be 10 C below ambient and an infrared camera was used to view the temperature gradient across the surface of the heat shield. At the transition region a clear temperature gradient will be visible, indicating whether the trip caused the flow to transition. Some results are shown in Figures 9 and 10. It should be noted that the contour range was adjusted for each run to illustrate the thermal variation and the location of the trip. Quantitative image-to-image comparison is not intended. The clean heat shield did not transition. The "=30, Re'=1000 0.0099 configuration transitioned for all angle of attacks consistently and was selected for the remainder of the test program. The distributing grit also transitioned near the stagnation point, but was not selected because of difficulty in application. The general findings from the study were:

! Trip sizes given by empirical method generally caused transition at Re' ~1000 location

! Re' 1000 location trip heights lower than indicated still caused transition.

! Re' 300 location trip heights from empirical method seem on the verge of tripping the boundary layer. They need to be higher for reliable tripping.

! Rek criterion was not a reliable metric for this subsonic / geometric regime.

Measurement of CM axial, tangential, and roll moment is shown in Figures 11 and 12 versus angle of attack. The axial force (free stream) and tangential coefficient (cross stream) were lower for the configurations that tripped the BL on the heat shield. No discernible trend was seen for the tangential force and roll moment for the different trip configurations.

"=0!, Re'=300

"=30!, Re'=300

"=30!, Re'=1000

Figure 8. Red lines are the """"=30 and Re''''=1000 trip dot pattern selected for the test. The gold lines were also investigated, and shown for reference.

Figure 9. IR thermography images of the smooth heatshield at """"=0 (left) and 30 (right). The flow does not trip on the heat shield.

Figure 10. """"=30 and Re''''=1000 IR thermography images of the turbulent transition at the trip dot location for 20 and 88 m/s. The flow trips at both speeds at the dot location.

Figure 11. CM axial force coefficient versus angle of attack. Legend indicates """"design/Re''''/Rek.

Figure 12. CM tangential force coefficient versus angle of attack. Legend indicates """"design/Re''''/Rek.

7Drag Performance

Load cell data were averaged for each run and vectorally summed to obtain the total parachute drag. Drag coefficient was defined using the following equation where Q is the dynamic pressure of the free-stream [8].

QSDrag

QD

DragCd02

04

%% $ (1)

The unreefed nominal configuration drag coefficient is shown in Figure 13. Drag appears to be maximum at "=30 and minimum at "=50 (capsule angle of attack). The drag coefficient at x/d=6 varied from 0.55 to 0.59. Full scale drogue parachute drag in the wake of a stream-lined body has previously been measured to be 0.61[1]. It is expected that the wake deficit would result in reduced drag performance relative to an unperturbed free-stream inlet condition. The dynamic behavior of the drag was evident in the high speed load trace (Figure 14). Inspection of the high speed video indicates a minor breathing mode in the skirt. The nominal inflated shape is shown in Figure 15 for "=0 and "=30. The inflated shape is not sensitive to the capsule angle of attack. Preliminary comparison to the full-scale profile indicates a representative shape was obtained with the subscale articles.

Figure 13. Unreefed or full-open parachute drag coefficient (Cd) as a function of capsule angle of attack at x/d=6. The drag coefficient is maximum at a CM angle of """"=30 and minimum at """"=50, for the angles measured.

Figure 14. Example of high speed force data for unreefed drogue at x/d=6 and """"=0. The largest RMS spikes are from the parachute tension force.

Figure 15. Inflated shape or profile of the unreefed drogue flying in the wake of the command model at (top) 0 and (bottom) 30 relative to the free stream.

Trailing distance

The nominal trailing distance for CPAS is 6d downstream of the CM. The trailing distance is measured from the CM maximum diameter. The experiment explored trailing distances at x/d=6, 8 and 9.5, to see if the parachute drag performance changed with increasing distance from the wake source. Figure 17 is a plot of the unreefed drag performance versus trailing distance for a "=0. Drag was found to increase with increased distance. To first order, inflated shape was not affected by trailing distance. A comparison of the inflated shape is shown in Figure 16 at x/d=6 and 9.5.

8Figure 16. Comparison of inflated shape of the unreefed parachute at (left) x/d=6 and (right) x/d=9.5

Figure 17. Unreefed parachute drag coefficient as a function of trailing distance at """"=0.

Reefed Performance

Figure 18. Comparison of inflated profile of the (top) 1stand (bottom) 2nd stage reefing at x/d=6, profile view. The 1st stage reefing has a smaller and more circular profile than the 2nd stage reefed. Both exhibit folds in the skirt at the reefing line attachment points.

As described previously, reefing is a means of controlling parachute load by reducing the open area of the parachute (with a reefing line around the skirt). Two-stage reefing is used with the CPAS drogue and was also explored in the test program. Figure 20 shows the inflated shape of the first stage and second stage reefing ratios. The parachute tended to fly more stably in the reefed state. The parachutes were never disreefed in the chamber. The dynamic motion of the parachute centerline relative to the capsule centerline was reduced. The reefed drag performance is shown in Figures 19. Cd,r is calculated relative to the full open area (So) to illustrate the reduction in drag from the reefing.

QSDrag

QD

DragC rd02

0

,

4

%% $

The oscillatory motion of the reefed parachute, in terms of trim angle and dispersion from the centerline, was reduced for the reefed configurations. This is likely due to a reduced tangential force coefficient, preventing a large trimming moment. The fabric dynamics were significantly increased, likely due to the more complex geometry of the reefed canopy opening to the in-flow environment. There was significant band flutter and projected area variation for the 1st and 2nd stage reefing.

Figure 19. First reefing stage drag performance.

Figure 20. Comparison of inflated canopy interior of the (left) 1st and (right) 2nd stage reefing at x/d=6.

Particle Image Velocimetry PIV measurements made upstream of the model parachutes

9turned out to be problematic. Laser light scattered by the suspension lines saturated portions of the PIV cameras charge-coupled-device (CCD) arrays, resulting in pixel saturation, as can be seen in the raw image in Figure 21. Since the saturation could potentially damage the CCD arrays, several attempts were made to mitigate the effect including painting the suspension lines black, but the final decision was to constrain the parachute laterally to the center of the tunnel by placing a rod through its vent and to image fields within the cone of the suspension line array (see Fig 6). This strategy, which was used successfully in a prior test program [13], ensured that scattering from suspension lines occurred only on the edges of the images. The image sets were also post-processed to eliminate erroneous vectors in any locations of strong scattering. Three runs were made in the constrained configuration. Results from two runs are shown in Figure 22. These include the mean streamwise (U) vertical (V), and spanwise (W) velocities as shown in Figure 6. U! refers to the free stream velocity and d the CM diameter.

The images are centered 0.1d above the wind tunnel and parachute centerlines and at the centerlines for the capsule only and parachute runs, respectively. In the constrained chute case, the streamwise velocity is decreasing as the flow approaches the canopy mouth. In the capsule only case, the field is very uniform over the field (to within 0.5%) with a streamwise velocity of 0.97U". In the parachute case, the streamwise velocity decreases from 0.92U" to 0.88U"across the field of view. The vertical velocity component is slightly upward across most of each plane. In the capsule

only case, the value decreases from 0.013U" to 0.009U"moving upward across the plane. In the parachute case, the value is nearly zero at the bottom of the plane and increases to 0.015U" near the top. The mean spanwise velocity is very small throughout the field (< 0.005U") in both cases as expected from symmetry. These datasets and others at angle of attack will be compared to CFD simulations of comparable flow fields at a later date. .

PIV measurements were also made at several axial locations downstream of the CM, to measure the wake recovery without the parachute present. Measurements were made at x/d=3.5, 5.5, and 7.1 for 0, 30 and 50 angle of attack of the CM relative to the free-stream. Data reduction is underway.

Figure 21. PIV camera image of the reflection of light off of the suspension lines. These represent regions of pixel saturation on the CCD array.

4. FUTURE WORKThe wind tunnel test program is still underway. An additional entry is planned to make measurements of internal pressure distribution with miniature pressure transducers on the canopy interior, as shown in Figure 23.

Figure 22. Contour plots of the PIV 3-component velocity measurement of the flow field at x/d =8 and """"=0. The top row of images is for the capsule only run with the field of view (FOV) centered 50 mm=0.1d above the

centerline. The bottom row is with the parachute present with the FOV at the centerline and 0.5d upstream of the parachute opening. The flow is from left to right with U!= 88 m/s and Q = 4788 Pa (100 psf) for both data sets.

10

Also measurement of parachute performance without the CM (reduced wake) will be conducted (Fig. 24). RMS load data will be analyzed to determine frequency content and RMS value. The high speed video data will be analyzed to determine trim angle, coning frequency, and projected area variation.

Figure 23. Planned locations of the ten pressure transducers on the canopy interior as indicate by the orange dots. Each sensor will be sewn to a radial.

Figure 24. No CM present parachute to wind tunnel mount configuration. The configuration is intended to simulate a no wake condition.

5. CONCLUSIONSA 10% of full scale wind tunnel test program to measure the Orion drogue parachute performance is underway. The program has measured spatially and temporally resolved velocity field upstream of the parachute with PIV, canopy dynamics with high speed video, and time resolved canopy and capsule forces with a force balance and 3-axis load cell. Reefed and unreefed configurations were investigated.

Results to date indicate similar performance and inflated shape to the full scale parachute with a reduction in drag performance due to the presence of the CM wake. Capsule angle of attack and trailing distance were also varied allowing quantification of drag and dynamic stability over a previously unexplored parameter space.

7. ACKNOWLEDGEMENTS The authors would like to acknowledge the staff at the Oran W. Nicks Low-Speed Wind Tunnel at Texas A&M

University and Phil Stuart, Allen Schwing, and Derek Dinzl for computational fluid dynamics expertise and support. This work was carried out at the Jet Propulsion Laboratory, California Institute of Technology, under a contract with the National Aeronautics and Space Administration.

REFERENCES [1] Randy Olmstead et al., Overview of the Crew

Exploration Vehicle Parachute Assembly System (CPAS) Generation I Drogue and Pilot Development Test Results, AIAA-2009-2939.

[2] D. Lichodziejewski et al., Development and Testing of the Orion CEV Parachute Assembly System (CPAS), AIAA-2009-2938.

[3] A. Sengupta, R. Machin, G. Bourland, R. Sinclair, P. Stuart, E. Hennings, E. Longmire, and A. Schwing, Subscale Test Program for the Orion Conical Ribbon Drogue Parachute, 21st Aerodynamic Decelerator System Conference, Dublin, Ireland, May 2011.

[4] D.E. Reichenau,. Aerodynamic Characteristics of Disk-Gap-Band Parachutes in the Wake of Viking Entry Forebodies at Mach Numbers from 0.2 to 2.6, AEDCTR72-78, 1972.

[5] J. Barber and H. Johari, Experimental Investigation of Personnel Parachute Designs Using Scale Model Wind Tunnel Testing, AIAA 2001-2074.

[6] J. Cruz et al., Wind Tunnel Testing of Various Disk-Gap-Band Parachutes, AIAA 2003-2129.

[7] A. Sengupta et al., Results from the Mars Science Laboratory Parachute Decelerator System Supersonic Qualification Program, Proc. of the IEEE 2008 Aero Conference, Big Sky, MT, March 2008.

[8] T.W. Knacke, Parachute Recovery Systems Design Manual, Para Publishing, Santa Barbara, CA, 1992.

[9] http://www.elsensors.com/html/kulite/LQ.LE-062.pdf, Last viewed November 1st 2011.

[10] Daniel C. Reda, Review and Synthesis of Roughness-Dominated Transition Correlations for Reentry Applications, Journal of Spacecraft and Rockets, Vol. 39, No. 2, MarchApril 2002.

[11] D. C. Reda, M. C. Wilder, D. W. Bogdanoff, and D. K. Prabhu, Transition Experiments on Blunt Bodies with Distributed Roughness in Hypersonic Free Flight, Journal of Spacecraft and Rockets, Vol. 45, No. 2, MarchApril 2008.

[12] D. C. Reda, M. C. Wilder and D. K. Prabhu, Transition Experiments on Blunt Bodies with Isolated Roughness Elements in Hypersonic Flight, Journal of Spacecraft and Rockets, Vol. 47, No. 5, SeptemberOctober 2010.

[13] A. Sengupta , J. Roeder, R. Kelsch, M. Wernet, A. Witkowski, and M. Kandis, Supersonic Performance of Disk Gap Band Parachutes Constrained to a 0 Degree Trim Angle Journal of Spacecraft and Rockets, Vol 46, No. 6, Nov 2009.

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BIOGRAPHIES

Dr. Anita Sengupta is a Senior Systems Engineer at NASAs Jet Propulsion Laboratory working primarily on entry missions and related technology developments for Mars, Venus, and Earth return. She received her PhD and MS in Aerospace Engineering from the University of Southern California and

BS in Aerospace Engineering from Boston University.

Dr. Ellen Longmire received an A.B. in Physics from Princeton University and MS and PhD in mechanical engineering from Stanford University. Since 1990 has been a professor in the Department of Aerospace Engineering at the University of Minnesota.

Mr. Mitchell Ryan is a recent MS graduate from the University of Minnesota in Aerospace Engineering. He is currently a Research Fellow in the Aerospace Engineering Department.

Mr. Erik Haugen is a graduate student in Aerospace Engineering from the University of Minnesota He also received his BS in Aerospace Engineering from the University of Minnesota in Aerospace Engineering.

Mr. Jose Laguna is a design engineer at Airborne Systems. He is a Member of the Orion parachute system design and test team. He received his BS in Aerospace Engineering from California State University at Long Beach where he is currently pursuing his MS in Aerospace Engineering.

Mr. Robert Sinclair is the chief engineer of Airborne Systems leading the the Earth Landing System development for the Orion Vehicle. He has been working in the field of parachutes for over three decades.

Dr. Edward White is an associate professor in the Department of Aerospace Engineering at Texas A&M University. He received his PhD in aerospace engineering from Arizona State University and his MS and MS from Case Western Reserve University.

Mr. Ricardo Machin is the chief engineer of the Orion Parachute System at the NASA Johnson Space Center. He received his BS degree in Aerospace Engineering from the University of Alabama.

Mr. Gary Bourland is an Aerospace Technologist at the NASA Johnson Space center, involved with parachute trajectory simulations, and subsonic and hypersonic wind tunnel tests. He has a BS in Petroleum and Aerospace Engineering from the University of Texas at Austin.

Ms. Elsa Hennings is the Chief Engineer for the Warfighter Systems and Support Division at the Naval Air Warfare Center Weapons Division. She currently serves as a subject matter expert for NASAs Orion Parachute system. She is a NAVAIR Research and Engineering Fellow, and received her BS in Mechanical

Engineering from the University of Missouri Columbia.

Dr. James Ross has worked at NASA Ames since 1980 in a variety of positions. He is currently the Coordinator of Aerodynamic Testing for the Multipurpose Crew Vehicle Aerosciences Project. He has a B.S. in Mechanical Engineering from UCLA, a Master of Engineering from California Polytechnic State University in San

Luis Obispo, and a Ph.D. from Iowa State University.

Mr. Daniel Bissell is an Application Engineer with TSI Incorporated in the experimental fluid mechanics division. Dan earned his Bachelor's of Aerospace Engineering and Mechanics (BAEM) at the University of Minnesota and is completing his Master's Degree in Mechanical Engineering at the University of St. Thomas.