Mars Precision Landing Using Guided Parachutes John B. McKinney 1 The Boeing Company, Huntington Beach, CA, 92647

Transcription

1 2th AIAA Aerodynamic Decelerator Systems Technology Conference and Seminar<BR> 4-7 May 29, Seattle, Washington AIAA Mars Precision Landing Using Guided Parachutes John B. McKinney 1 The Boeing Company, Huntington Beach, CA, Charles H. Lowry 2 Consultant, Garden Grove, CA, A method for precision landing of manned spacecraft and cargo modules on Mars using guided parachutes was successfully demonstrated in terrestrial tests. A simulation model was developed based on those tests to explore the full capability of this type of system. The Mars Guided Parachute (MGP) concept of minimizing Mars landing site errors was found to be a practical and valuable approach. For this effort, off-the-shelf hardware was selected and assembled. Aerial drop tests were conducted on Earth at Mars-relevant atmospheric densities to collect representative performance data on how well such a system would traverse upon command to a desired landing target. A 3 degree of freedom (3 DoF) model was employed to evaluate a Mars-scaled landing system based on conservative inputs from the terrestrial tests. A wide range of parameters including system design characteristics and wind and navigational errors were perused to explore MGP system capability. Results were extremely favorable. I. Introduction The objective of the Mars Guided Parachute (MPG) program was to demonstrate the technical feasibility of reducing landing site errors on Mars during descent by using parachute Wind Drift Compensation (WDC) control. The technology developed for JPL would be used during Entry, Descent, and Landing (EDL). This work supported the NASA NRA 3-OSS-1, Amendment A.2.2, Mars Exploration Program Advanced Technologies. Specifically, it addressed Technology Area 2.5, Advanced Entry, Descent, and Landing, and is one component of a set of comprehensive EDL system technologies that support pinpoint landing to within tens of meters of a target landing site. This technology applies to both manned and unmanned landings on Mars. As robotic and eventually human exploration of Mars continues to expand, the ability to place payloads on the surface of Mars with increased precision is required. Science packages have begun to investigate particular sites and features. Future robotic missions may combine the capabilities of multiple entry systems requiring close proximity landing. For instance, a sample return mission might use separate entry vehicles for the sample gathering and sample return capabilities. Human exploration will likely require crewed entry vehicles to land in close proximity to habitats and supplies. For these reasons the Mars Technology Program is working on greatly reducing the entry and landing errors throughout the EDL system. Other programs are investigating controlling errors during hypersonic entry, through lift vector rotation. Similarly, other teams are investigating methods for enhanced precision in the final moments of landing, including additional features to provide a level of hazard avoidance. Parachute flight, while brief, remains a vital portion of the EDL sequence at Mars and other planets with a significant atmosphere, at least for most re-entry vehicles. Therefore, a level of control during this portion of the descent phase can provide a significant improvement in landing precision. Guidance, Navigation and Control (GN&C) can compensate for significant errors in knowledge of atmospheric variables such as wind velocity and atmospheric density. To a lesser extent, GN&C can also compensate for errors in the initial condition for the parachute descent phase due to similar atmospheric errors in the previous EDL phase, such as the heat shield entry flight. 1 Space Technology Research & Development Engineer, Dynamics and Smart Structures, 149 Bolsa Chica, H13- C326 Huntington Beach, Ca 92647, AIAA Associate Member 2 Parachute Consultant, 932 Crosby Ave, Garden Grove, CA, 92844, AIAA Senior Member 1 Copyright 29 by the, Inc. The U.S. Government has a royalty-free license to exercise all rights under the copyright claimed herein for Governmental purposes. All other rights are reserved by the copyright owner.

2 A number of earthborne guided parachute systems are either available or being developed for a wide range of military cargo applications. These cover a range of weights from tens of pounds to tens of thousands of pounds. The current approaches include a variety of parachute systems from guidance of traditional cargo class parachutes, parafoils for very large cargo systems and high speed flight of the cargo transitioning to a last minute, un-controlled parachute landing. In the Mars environment, there have been only a handful of missions and perhaps 2-4 handfuls of development tests spanning the past 4 years. The reason for this is directly related to the cost of the testing. The best earthborne test of a Mars relevant system requires delivery of the payload to an altitude of approximately 35 KM (115 K ft.), typically by balloon. However, if the correct entry velocity is also required, depending on the parachute system, a rocket acceleration stage may also be required. For full system demonstration the correct shape of the entry vehicle may be required, further complicating (and raising the expense) of the testing. The MGP program had much more modest goals. Our primary mission was to: 1. Demonstrate through test that a parachute can be guided in a Mars- relevant environment. 2. Demonstrate through simulation the value of such a control feature. This program was initiated August 19, 24 by JPL Contract P.O The program involved a Boeing led team including Irvin Aerospace, Global Solutions for Science and Learning (GSSL) (now NSC), Vertigo, and Oregon State University. The research has been accomplished by means of four principal tasks. First, a computer simulation was developed to provide a means to predict controlled descent performance in various Martian wind environments. Second, an open-loop high altitude flight test was conducted in the 1 st program year to validate parachute aerodynamics at an equivalent Martian atmospheric density range. Third, a closed-loop flight article was designed, assembled, and integrated in the 2 nd year. Finally, a closed-loop flight test was planned for the 3 rd year and conducted to demonstrate fully autonomous controlled descent and to provide inputs to the simulations that followed. Results of some aspects of the program are presented in this paper. Boeing Report PWDM7-8, Mars Guided Parachute (MGP) Design and Flight Test, Final Report, dated April 3, 27 documents this total effort. II. Formulation of the problem As robotic exploration of Mars continues, science objectives are becoming increasingly focused on investigation of specific locations and features on the planet. This in turn results in increased focus on the ability to place a lander and instruments in particular locations. Current technology landing dispersion ellipses are much larger than the requirements of future missions. The Advanced EDL Systems segment of MTP deals primarily with approaches to improve precision in landing point guidance. During the parachute portion of the EDL event, the spacecraft is subject to drifting with current winds aloft. Therefore a parachute that can be controlled through the generation of a lift vector could correct for errors in these predicted wind profiles and possibly also provide some correction to other entry dispersions. This approach is currently being used by the US Army in a number of cargo delivery systems. The most relevant being the Affordable Guided Airdrop System (AGAS) from Vertigo. AGA utilizes a round parachute that can be slipped into a desired glide direction without need to turn the parachute. This results in a much more responsive system. An MGP program goal was to demonstrate that an AGAS-like approach to parachute control is viable in the Mars relevant flight environment. We adopted the AGAS approach as the development team was quickly convinced that an investigation of parafoil inflation and flight would face significant technical challenges. There are significant differences, of course, in using AGAS in Earth s atmosphere vs. on Mars. The atmospheric density differences between Mars and low altitude Earth (less than 2 K ft.) create significant differences in the parachute systems. The thin atmosphere at Mars creates significant mass differences in the systems. In the Mars case, the parachute/cargo system center of mass and center of gravity are very close together. Mars missions generally operate at significantly higher airspeeds than low altitude terrestrial systems. This will and has been shown to produce a parachute system that responds faster to control inputs than the terrestrial cargo delivery systems. Finally, the low density of the Mars relevant environment, including high altitude at Earth, provides significantly 2

3 less damping than the low altitude Earth environment. This has been shown to lead to oscillatory flight conditions in other systems tested in the Mars relevant (High Altitude Earth) environment. Our approach to demonstrating the MGP application involved high altitude flight testing on Earth. Flight at altitudes in the vicinity of 35 KM (115 K ft.) provides atmospheric density similar to the Mars environment. As we are interested in the flying characteristics of the parachute, not the inflation characteristics and loads, acceleration of the payload to flight speed and shape modeling of the spacecraft forebody are details that were not required. Thus, the MGP program could test in a Mars relevant environment through relatively simple balloon launches at the required altitude. This approach also provided for affordable flight opportunities consistent with the investment being made. However, balloon launch of payloads similar to future JPL missions to Mars, such as Mars Science Laboratory (MSL) becomes expensive due to the size of balloon required MSL would need to launch approximately 2 Kg. Thus, our approach was to test at reduced mass and relate these results back to future missions through a set of scaling laws. In the end, the 1 M diameter Beagle 2 Ringsail parachute was shown to scale effectively to the MSL mission and it represents a recent Mars mission at full scale. The AGAS approach divides the parachute rigging into 4 sets of lines, attached to one riser from each quadrant of the parachute. By adjusting these quadrants shortening or lengthening the risers, the parachute is made to glide in one of four (1,3,5,7) primary directions Figure 1. Glide Directions Similarly, adjustments to pairs of risers can cause glide in the four directions (2,4,6,8) between the four primary directions. This approach produces a system that can glide in a total of 8 directions (Figure 1). This, combined with knowledge of which direction the system needs to glide, creates an effective control system. Figure 2 presents a view of the deformed parachute in glide mode during flight testing. 3

4 Figure 2. Beagle 2 in Gliding Flight III. Balloon Drop Tests The main thrust of the program included two balloon launches to demonstrate the concept at Mars-relevant conditions, and simulations to demonstrate its value to EDL. A simple low altitude flight test program was added early in the program to confirm glide obtained with the Ringsail class parachute. Flight tests were conducted on an opportunity basis at a local drop zone. Ten M Ringsail parachutes, with an appropriate payload were deployed at low speed and shortly after inflation, pre-timed cutters released the riser legs and introduced the glide command. Instrumentation on the test parachute included GPS position and pressure altitude. Non-gliding Tri-lobe parachutes with similar instrumentation were deployed at the same time to provide a reference for winds at altitude. Tri-Lobe Beagle II Figure 3. Early Test GPS Data 4

5 The difference between the two position time histories was used to remove the effect of air mass movement (winds) and the parachute glide ratio was computed. Figure 3 provides an example of the GPS data gathered. Further reduction of this data indicated an average glide ratio of slightly greater that.7, which was consistent with glide ratio developed for previous AGAS systems. It was therefore concluded that no parachute modifications were required. The purpose of the first balloon (Open Loop) flight was to obtain parachute glide and control response characteristics to support simulation and assessment of the flight control laws. The second balloon test (with 2 gliding parachutes set at 2 different glide ratios) demonstrated closed loop control of the parachute: both tests were in the Mars relevant environment. The simulations covered in later paragraphs then took the system characteristics as determined by the drop tests and produced a broad validation of the MGP concept. These major test/simulation events are discussed in the following paragraphs. Figure 4. Test Payloads The first balloon test (Open Loop) provided boundary case data on glide ratio for this chute at high altitude. After approximately 7 s of well behaved glide and data acquisition, it entered a spin which persisted down to lower altitudes. It was determined that overdriving this Beagle chute to such high glide ratios (~.7) was not wise, so more moderate glide ratios were commanded for the second (Closed Loop) balloon test. The second balloon test (Figures 4 and 5) was the key event of the program. Among the details to be measured in this test were the actual glide ratio and total response time for a given control input as well as overall glide performance for the 2 gliding parachutes. 5

6 The two gliding parachutes in this drop were designated as MGP 1 and MGP 2, respectively. Also, the actuation lengths (defined as riser stroke as percentage of canopy reference diameter, D o ) for MGP 1 and MGP 2 were ± 2% and ± 3% D o, respectively. It was predicted that these riser strokes would produce glide ratios of.2 to.25 and.4 to.45 respectively. (This can be compared to the Open Loop Drop Test that glided the parachute at an actuation length of ± 6% D o ). Figure 5. 2 nd Balloon Drop In addition to the two gliding parachutes, the balloon carried aloft a non-gliding Tare parachute to descend in real-time with the test units and record wind drift. Primary instrumentation consisted of GPS on all three parachute payloads and a magnetic compass on the two gliding parachutes. On Sept 26, 26, the MGP systems suspended under a high altitude balloon, began ascending to the drop test altitude. Weather conditions were ideal with no clouds and light winds. The balloon ascended for two hours to a float altitude of 34.7 km, and after 29 min at float altitude, the Tare parachute was released at 11:24:48. MGP 1 was then released at an altitude of 34.7 km, after the Tare parachute was dropped, at 11:26:25. After the MGP 1 was dropped, the MGP 2 flight unit was released at 11:33:34. Without the ballast weight of the MGP 1 unit, the altitude of the balloon increased to approximately 36.4 km prior to releasing MGP 2. On-board video and telemetry data confirmed all parachutes deployed and inflated properly. Each gliding parachute was commanded to fly due West for 9 s, followed by East for 9 s, and finally back to the West for another 9 s. The Tare parachute and the MGP 1 unit functioned as expected during the test. However, the MGP 2 flight instrumentation system exhibited two flight anomalies: A fault in the video system resulted in a loss of in-flight video, and an issue in the GPS transponder or data acquisition system caused sporadic behavior in the recorded data trace. Although the MGP 2 flight data appeared problematic, satisfactory gliding performance was observed in the system. After reducing the test data, the MGP 1 and MGP 2 units were shown to demonstrate the intended controlled flight behavior in the relevant Mars atmosphere. 6

7 Detailed in the following paragraphs are the results of the test as well as the data reduction performed to characterize the overall system performance. Throughout this document, the cross range trajectory plots are designated as being either wind or no wind graphs. The wind designation delineates a plot of the payload with the bias of the wind. The no wind designation outlines the trajectory that would have been adopted by the payload in the absence of wind. Figure 6 shows the trajectory adopted by the MGP 1 unit. In the graph, a distinct West, followed by East and subsequent West trajectories are seen. This result indicates that the controller is augmenting the flight of the parachute to accommodate the desired trajectory. Figure 7 outlines the east track versus time. No Wind Trajectory t = 2 s West Altitude (m) t = 11 s East t = 2 s West East (m) Figure 6. MGP 1 Altitude vs. East Trajectory with Glide Direction Command Points t = 2 s West t = 2 s West -15 East (m) No Wind Trajectory t = 11 s East Time (s) Figure 7. MGP 1 East Trajectory vs. Time 7

8 A significant change in heading during the flight was observed in the upward facing video of the flight. Figure 8 shows the net heading change based on the location of the sun with respect to the parachute canopy in the flight video. The slipping of the canopy toward the desired glide direction constantly compensated for this unwanted heading change Total Rotation (deg) Time (s) Figure 8. MGP 1 Observed Heading Change. The glide ratio for this drop ranged from approximately.25 to.45 during the time frame of interest. Note that the glide ratio in the figure includes the vector sum of cross range velocities in both the East and North directions, but the North component was small. The ± 2% commanded riser actuation stroke was designed to provide a glide ratio of.2 to.25, so it appears the desired glide ratio was exceeded on this drop. It was also noted from the data that the glide ratio increased as altitude decreased. Looking now at the MGP 2 reduced flight data, the following East ground track versus altitude was constructed, Figure 9. Figure 1 shows the time history of the East track. 8

9 No Wind Trajectory t = 2 s West t = 11 s East Altitude (m) t = 2 s West East (m) Figure 9. MGP 2 Altitude vs. East Trajectory with Glide Direction Command Points t = 2 s West -3-4 East (m) -5-6 t = 2 s West -7-8 t = 11 s East -9-1 No Wind Trajectory Time (s) Figure 1. MGP 2 East Trajectory vs. Time 9

10 The magnetic heading of the raw MGP 2 information was similar to MGP 1 but accurate glide ratio data was not obtained due to the aforementioned instrumentation problems. The ± 3% riser actuation stroke was designed to provide a glide ratio of.4 to.45 and these values were not apparent in MGP 2. A possible cause is scatter in the data which has been attributed to the issues in the MGP 2 instrumentation system. IV. Conclusions from Closed Loop Drop Test The results of the test indicate that both systems demonstrated the ability to control the trajectory of the payload. The feasibility of coupling a closed loop control system and a round parachute to achieve controllable glide was successfully demonstrated. The tendency of the 1 m Beagle 2 Ringsail parachute canopy to enter a spiral flight path at these high altitudes was avoided by commanding to more moderate glide ratios than were used on the Open Loop Drop Test. Significant parachute pitch, roll, and yaw were evident at these high altitudes. The excessive yaw rates caused the control system to continually seek to correct the directional flight path. Since, by design, it could only accept a GPS-directed command every 7 s, the parachute tended to wander off course significantly before the next correction could be executed. However, the overall trajectories were very much along the commanded directions. Any effort to further develop this type of system would benefit from employing a more stable parachute canopy for this high altitude environment, even if it sacrifices drag performance. Glide response-time to commands was measured, but the above stability issues prevented the recording of consistent data on this test. The response data from the Open Loop drop test provided good data characteristic of this Beagle 2 parachute, where the glide ratio of approximately 4.5 s was established while little yaw was occurring. Glide ratio values for the two parachutes were determined. The resulting glide ratio for MGP 1 was in the range of approximately.25 to.45 (as compared with the pre-test desired value of.2 to.25). The glide ratio range determined for MGP 2 was similar to that of MGP 1 (as compared with the desired.4 to.45). Data issues may have affected the MGP 2 values, so the MGP 1 data is considered the most reliable glide ratio values V. Simulation In conjunction with the conducted tests, the primary goal of the MGP simulation effort was to demonstrate the technical feasibility and value of reducing Mars landing site position errors by using a WDC algorithm. To accomplish this goal, a model was developed to investigate and evaluate the important parameters in minimizing landing position errors. This model considers wind prediction errors, navigational errors, and various MGP system design characteristics. The simulation runs considered a realistic size terminal descent parachute and payload weight for future Mars landings. The MGP model used in this effort is a three translational degrees-of-freedom dynamics model with lift and drag forces and apparent mass corrections. Rotational equations of motion are not considered, which implies a stable parachute. But inputs to the model (discussed later) do acknowledge the chute instabilities experienced in the drop tests. This model incorporates an aerodynamic model of the parachute and a functional model of the control system. The Oregon State University Mars wind models developed for this program were used for wind input. The MGP model determines a reference trajectory by back-propagating the estimated lateral wind velocity and terminal descent velocity emanating from the desired landing spot and ending at some point in space at a desired altitude. This point in space is where the parachute should be deployed so as to land at the desired spot on the ground in the presence of a wind. It then performs gliding maneuvers to compensate for both navigation and wind errors encountered. The model has been enhanced by adding hysteresis compensation for reduced final landing position error and to decrease the oscillations of the actual trajectory about the reference trajectory. 1

11 The MGP model emulates the flight controller, or AGU, which functions as follows. The system works through prior knowledge of the winds in the deployment area. A reference trajectory based on the system's descent rate is generated as stated earlier on a mission computer, and this trajectory is transmitted to the flight controller of the AGU. When the system is deployed, the AGU uses a navigational system to determine its location relative to the reference trajectory. If the distance to the reference trajectory is greater than two times the diameter of the inflated canopy, the system performs a balanced riser slip in the direction of the reference trajectory. The system continues to make steering corrections toward the reference trajectory until it is within 1.5 diameters of the reference trajectory. Once within 1.5 diameters, the system returns the risers to their neutral state, and they remain that way until the system lands or drifts out of the two diameters. Note that the test system performs trajectory corrections based solely on position deviations from the reference trajectory, and not on velocity. Figure 11 depicts how the system works. Figure 11. Precision Landing Mission Profile The flight controller was modeled to be identical to the AGAS unit used in the drop tests. Steering corrections are accomplished by slipping the parachute in any of 8 directions by actuating winches in the flight controller. Again, the parachute is not rotated, or yawed, to change glide direction rather the skirt is deformed by shortening/lengthening one or 2 (of 4 total in the chute) selected risers to glide the chute in the desired direction. Two winches at right-angles to each other provide north-south and east-west glide. Combined action of both winches provides glide in intermediate directions (such as northeast). Each winch can be driven in forward or reverse directions rotating a winch in one direction reels in (shortens) a riser, while extending the opposite riser. The parachute glides in the direction of the shortened riser. Once actuated, a winch goes full stroke there is no proportional control in this test system. 11

12 A matrix of inputs and conditions was run to explore MGP system capability and value in minimizing landing errors. Parameters were selected to represent what was considered nominal and off-limit to ensure that the scope of simulated conditions was sufficiently broad to be enlightening. Although the parachutes tested in this program were 1 m in diameter, the simulation runs comprised a chute size of 45 m and a payload weight of 2 kg. Certain assumptions related to spacecraft navigation were necessary to set up a starting point for the MGP simulations. The spacecraft enters the Mars atmosphere and navigates so as to attach to the reference trajectory at the predicted deployment point, at a reference altitude. As stated earlier, the trajectory is intended to lead ultimately to the target landing site. This reference trajectory considers the wind profile predicted for the time of landing, and if there were no errors in wind predictions or navigational errors, the spacecraft would land directly on the target without need for any gliding from the MGP. The directed gliding feature of the MGP is required to deal with the errors that will certainly exist thus the MGP and this simulation. In reality, the spacecraft enters, decelerates due to its own base drag, goes through a drogue stage to further slow it, and finally arrives in the neighborhood of the computed terminal descent chute deployment point, which is the point at which the simulation starts. At this point, most horizontal velocity has been taken out by drogue stage drag and the nominal vertical velocity may be close to the drogue terminal velocity (which varies primarily with atmospheric density variations). For purposes of the simulations, residual velocities are taken to be Mach.1 in X and Y directions and Mach.8 in the vertical (down) direction. From the deployment point, the simulation allows a 48 s period for inflation, stabilization, and deceleration prior to issuing a glide command. Navigation errors of, 5 m and 1 m (expressed in error circle radius) were selected and applied based on the state-of-the-art of such systems. The method of applying the error circle was always worst case the 5 or 1 m error was taken in a direction to place the spacecraft furthest away from the target landing site and the Mach.1 residual velocity directed away from the target. The navigational error could also be in the altitude, so runs were also conducted with initial altitude 5 m higher than the reference altitude hr 3hr 6hr 12hr 1 3 Altitude NormofWindEst. Err.(m/s) Figure 12. Wind Errors 12

13 The predicted winds were provided by Oregon State University models as discussed in prior paragraphs. The predicted wind profiles were developed for each hour of the day. Wind prediction errors for model input were simply derived by using wind profiles of 1, 3, 6, 9, and 12 hours before the planned landing. This time-offset provided light, nominal, and severe wind errors for a large range of simulation conditions. The predicted winds in a typical run may be in the vicinity of to 1 m/s as the spacecraft descends, but the 1 to 12 hr wind error increases it as shown in Figure 12, which is an RSS plot of OSU predictions from their North Polar Cap site data. The Open-Loop and Closed-Loop Drop Tests provided MGP system performance inputs to the simulations, particularly in the parachute area. Drop MGP 2 from the Closed-Loop test provided an average glide ratio value of.345 from the first 2 legs of the flight. The measurements leading to this value evaluated the slope of the Altitude vs. Horizontal Range curve, and it included the reality that the test chute was experiencing considerable yawing (and, no doubt, pitching and rolling) during the measurement time. This would imply that a more stable parachute (such as one developed specifically for Mars duty) would have a larger effective glide ratio, so an increased but achievable glide ratio value of.7 was also chosen as an input to the simulation. Further, the glide ratios predetermined for the Closed Loop test were chosen conservatively to avoid the spiraling experienced on the previous Open Loop test. This implies that a higher glide ratio could perhaps have been chosen had additional knowledge been available relative to the no-spiral threshold. Response time is the time for the chute to respond fully (in terms of achieved glide velocity) to a glide command. In the model, the winch reaches full stroke 2 sec after command. The chute begins to deform and glide as the winch starts to stroke and proceeds toward full glide velocity in a first-order exponential fashion. The chute response time for the Beagle 2 parachute was determined to be approximately 4.5 s. Analysis showed that a factor of 3.4 should be used to project from the test chute size, 1 m, to the 45 m chute. Thus the response time for the larger chute was determined to be sec. The flight controller has a preset deadband, which is the time between a given glide command and the point in time it will accept the next command. This deadband is necessary to allow the chute time to develop lift (glide) in response to a given command before being interrupted by a subsequent command. The deadband need not be so long as to allow full glide velocity to be achieved, and thus could be less than chute response time. But it could also be longer than the chute response time. So a range of possible deadbands were run, from 7 s (which was used in the drop tests) to 2 s. Deployment altitude was a key variable in the simulations. A range of deployment altitudes (from 4 to 15 km), spanning the experience of past missions to Mars, was used in the MGP simulations. The primary concern with any system seeking to provide precision landing capability is simply expressed by the question How close to the target can we land? The target, in this case could be a stake driven in the soil at the exact desired landing spot. The results of the MGP simulation runs, however, point out that there are really two concerns. One is long distance performance, or, What is the system s capability to reach the vicinity of the target from a distance? The other is close-in performance, or, What is the system s capability to come to rest on or near the target instead of limit cycling around it?. The aforementioned imaginary cylinders, specifically the outer one, is the target for the long distance trip from the point of terminal chute deployment. Navigation errors, wind errors, MGP glide ratio, and deployment altitude all influence the ability to reach the cylinders. If the cylinders are not reached, the landing site error is simply the distance the system falls short of the target. But if the reach was adequate and the cylinders entered, the landing error becomes somewhat random. That is because the MGP may traverse over the target, then reverse itself a number of times, continuing to overshoot the target until ground impact. The time at ground impact is independent of where the system is in its limit cycle, so the landing error is small (as low as 2.7 m in this series of runs), but random. It is also possible for the MGP to exit the cylinders after perhaps several limit cycles within that is when a large ground wind strikes and the system does not have enough altitude left to correct for it. This is particularly likely to happen when the wind used for generating the reference trajectory is a few hours stale. Selected results of the MGP simulations are presented in this paragraph in tabular and graphic form. Table 1 is a matrix of the formal runs conducted, showing the parameters used and landing site error information in terms of whether it reached the threshold cylinders, and landed within or out. For those runs that landed outside the cylinders, 13

14 the landing distance from the target is listed. Dual glide ratios of.7 and.345 were explored in run 1k to simulate the effects of proportional control. The navigational errors (nav errors) are applied radially as discussed earlier, and in all runs an additional up error of 5 m was applied. The right-hand column includes key words expressing the purpose of the various run-sets. The purpose of applying this wide range of error conditions (navigational and wind) was to tax the system s ability to achieve precision landings. It is seen from this chart that success in reaching the vicinity of the target is achieved in a wide range of nominal and off-nominal conditions and that the MGP system is truly effective in minimizing landing site errors. Specific observations are diverse and informative. Under all worst case conditions of navigation errors and under all wind errors up through 6 hr age, the system reached the threshold cylinders, after starting from a 1km deployment altitude. When the deployment altitude is reduced to 5-6 km, the ability to reach the cylinders is reduced, as evidenced by runs 1f-1i. Increasing the glide ratio to.7 increases the likelihood of success, as shown by run 1j. Runs 8a and 8b, with wind age of 3 and 6 hr respectively, show that the system may easily reach the target threshold cylinders only to be blown off-track by high ground winds just before landing and left with the dilemma that not enough altitude remains to allow a correction. This condition could be lessened or avoided by the addition of proportional control as evidenced by run 1k which employs a.7 glide ratio initially, reducing to.48 when the outer threshold cylinder is reached. A smaller controller deadband is moderately helpful as shown by runs 8a, 16c, and 17c. Although not shown in the table, the number of winch actuations varied from 7 to 33. The most winch actuations result from attaining and attaching to the reference trajectory early in the flight, which results in limit cycling within the threshold cylinders. Conversely, the least number of actuations result from barely reaching the threshold cylinders before touchdown, as in runs with extreme errors or low deployment altitude. [The flight controller is mechanized such that stroking forces (shortening one riser while extending the opposite one) are well balanced, resulting in low energy demand. Calculations show that on Mars with a 45 m chute, 2 km payload, 2 strokes assumed, and a margin of 1, the battery weight would be approximately 5 Earth pounds]. 14

15 Run No. Glide Dead Band-s Depl. Radial Wind Reach Land In Miss Distm Ratio Alt.-km Nav. Error-m Disp. hr Cyls Cyls Purpose (GR) (+5Up)?? y y - No Wind Error 2 y n y y y y - Wind Error 7 y y - With wind & nav error 8a 3 y n 99 8b 6 y n n n 116 1a 1 1 y y - 1b 7 y y - Depl Alt 1c 6 y y - 1d 5 y y - 1e 4 n n 293 1f 1 3 y y - 1g 7 y y - 1h 6 y y - 1i 5 n n 17 1j.7 y y - Hi GR 1k.7/.48 y y - Dual GR y n 19 Hi Errors n n y n 18 Hi GR 16a y y - Dead Band 16b 1 y y - 16c 3 y n 74 16d 6 y n a 12 1 y y - 17b 2 y n 72 17c 3 y n y y y y - Depl Alt 2 7 y y y y - Table 1. Matrix of Simulated Runs Performed VI. Simulation Conclusions The MGP model developed for this task is a valuable tool for exploring and evaluating the MGP system and the many important parameters of interest. The MGP system is shown to be an effective means of minimizing landing site errors through a wide range of conditions. The runs show there is no optimum deployment altitude but altitude is a major determinant in achieving sufficient range to reach the target vicinity. Optimizing the parachute design for increased glide ratio was shown to further enhance reach. Dual glide ratios, high initially and lower when the actual trajectory enters the cylinders around the reference trajectory, are effective in minimizing landing site errors. The runs point out how a future system could further reduce errors by incorporating velocity (in addition to position) intelligence in the control logic. 15

16 In addition to validating the MGP concept based on the use of off-the-shelf hardware, the program provided insights into how a next-generation research thrust could be carried out without excessive expenditures, while providing large technical benefits. For example, the Beagle 2 parachute selected for this program had been designed for Mars duty, but it was used in this program in a gliding configuration which seemed to introduce stability issues. Future work in this area could employ a chute with aerodynamic features that enhance stability, and produce an even higher glide ratio. The nonlinear control system tested in this contract only used position error to initiate trajectory corrections. As the system approached the target, it tended to overshoot, then reverse and overshoot perhaps multiple times. Even then, with no large ground wind spikes, the target miss distance seen for this condition was less than the 72 m cylinder radius in most cases, as seen on Table 1. But a proportional control system using velocity intelligence as well could be easily mechanized to highly optimize close-in performance for future research. Figures 13 and 14 show typical runs where the guided parachute system seeks out and enters the target cylinders and lands within the inner cylinder. Reference and Actual Trajectory Ref. Actual 1 8 altitude N-S: xi (meters) E-W: yi (meters) 8 Figure D Plot of System Performance from Deployment Point to Landing Spot: Actual and Reference. 16

17 N-S Actual & Ref (m) Ground Trace in the Inertial frame Actual Ref E-W Actual & Ref (m) Radial Position Dvn (m) Altitude (m) Time(sec) Figure 14. Typical Simulation Runs Radial Position Dvn (m) References 1 Taylor, T., McKinney, J.B., Brown, G., Mars Guided Parachute, a Status Report 25, AIAA 25, 2 Taylor, T., McKinney, J.B., Brown, G., Mars Guided Parachute, a Status Report 27, AIAA 27, 17