The Advanced Supersonic Parachute Inflation Research Experiment (ASPIRE) was a series of sounding rocket flights aimed at understanding the dynamics of supersonic parachutes that are used for Mars robotic applications. Although the 2012 Mars Science Laboratory (MSL) had a successful deployment of a supersonic parachute, postflight analysis and the results of the Low-Density Supersonic Decelerator project called into question some of the assumptions underlying MSL's parachute qualification program. The ASPIRE project was a key part of the risk-reduction program undertaken by Mars 2020, the follow-on to MSL. The first sounding rocket (SR01) flight of ASPIRE occurred near Wallops Island, Virginia on 4 October 2017, and demonstrated the successful deployment and inflation of an MSL build-to-print parachute in flight conditions similar to the 2012 MSL mission. ASPIRE SR02 and SR03 were successful follow-on flights on 31 March 2018 and 7 September 2018 that demonstrated the new, strengthened supersonic parachute designed for the Mars 2020 project. The SR02 and SR03 parachutes were targeted to 100 and 140% of the expected flight limit load for Mars 2020 to confirm new margins expected from the strengthened parachute. Prior to all flights, a multibody flight dynamics simulation was developed to predict the parachute dynamics and was used, in conjunction with other tools, to target Mars-relevant flight conditions. After each flight, the onboard data were used to reconstruct the flight trajectory and to validate the preflight dynamics simulation. Postflight analysis showed that all three tests achieved their targeted conditions and preflight modeling bounded the key performance metrics for the parachute. This paper describes the flight mechanics simulation, postflight reconstruction, and the reconciliation process used to validate the flight models.

The Advanced Supersonic Parachute Inflation Research and Experiment (ASPIRE) is a series of sounding rocket flights aimed at understanding the dynamics of supersonic parachutes that are used for Mars robotic applications. SR01 was the first sounding rocket flight of ASPIRE that occurred off the coast of Wallops Island, VA on Oct. 4, 2017 and showed the successful deployment and inflation of a Mars Science Laboratory built-to- print parachute in flight conditions similar to the 2012 Mars Science Laboratory (MSL) mission. SR02 was the second sounding rocket flight that also occurred off the coast of Wallops Island on March 31, 2018 and showcased the successful deployment and inflation of a new strengthened parachute being considered for the Mars 2020 mission at fifty percent higher dynamic pressure than observed on MSL. Prior to both flights, a multi-body flight dynamics simulation was developed to predict the parachute dynamics and was used, in conjunction with other tools, to target Mars-relevant flight conditions. After each flight, the reconstructed trajectory was used to validate the pre-flight dynamics simulation and recommend changes to improve predictions for future flights planned for the ASPIRE pro- gram. This paper describes the flight mechanics simulation and the post flight reconciliation process used to validate the flight models.

The Low-Density Supersonic Decelerator (LDSD) project performed the second test flight of a supersonic inflatable device (SIAD) and ring-sail parachute on June 8, 2015 splashing down in the Pacific Ocean west of Kauai. In order for recovery ships to quickly extract the test vehicle hardware from the ocean, and in the interest of safety to the population of the islands of Kauai and Nihau, statistical estimates of the splashdown location had to be performed. This paper describes the modeling assumptions used to generate the Monte Carlo splashdown trajectory simulation results and the use of the splashdown probability ellipses in support of satisfying range safety requirements.

Increasingly complex space missions require powerful and flexible simulation environments in order to design, analyze, and operate the missions. NASA’s Jet Propulsion Laboratory has created the DSENDS simulation environment that addresses these needs for a wide range of space missions. In this paper, we describe the DSENDS simulation environment and the key architectural components that make DSENDS a useful simulation and analysis framework. We also overview a variety of NASA missions and flight experiments that are using DSENDS.

NASA's Low Density Supersonic Decelerator (LDSD) project conducted its second Supersonic Flight Dynamics Test (SFDT-2) on June 8, 2015. The Program to Optimize Simulated Trajectories II (POST2) was one of the flight dynamics tools used to simulate and predict the flight performance and was a major tool used in the post-flight assessment of the flight trajectory. This paper compares the simulation predictions with the reconstructed trajectory. Additionally, off-nominal conditions seen during flight are modeled in the simulation to reconcile the predictions with flight data. These analyses are beneficial to characterize the results of the flight test and to improve the simulation and targeting of the subsequent LDSD flights.

Program to Optimize Simulated Trajectories II (POST2) was utilized to develop trajectory simulations characterizing all flight phases from drop to splashdown for the Low-Density Supersonic Decelerator (LDSD) project's first and second Supersonic Flight Dynamics Tests (SFDT-1 and SFDT-2) which took place June 28, 2014 and June 8, 2015, respectively. This paper describes the modeling improvements incorporated into the LDSD POST2 simulations since SFDT-1 and presents how these modeling updates affected the predicted SFDT-2 performance and sensitivity to the mission design. The POST2 simulation flight dynamics support during the SFDT-2 launch, operations, and recovery is also provided.

The Low-Density Supersonic Decelerator (LDSD) Project's first Supersonic Flight Dynamics Test (SFDT-1) occurred June 28, 2014. Program to Optimize Simulated Trajectories II (POST2) was utilized to develop trajectory simulations characterizing all SFDT-1 flight phases from drop to splashdown. These POST2 simulations were used to validate the targeting parameters developed for SFDT-1, predict performance and understand the sensitivity of the vehicle and nominal mission designs, and to support flight test operations with trajectory performance and splashdown location predictions for vehicle recovery. This paper provides an overview of the POST2 simulations developed for LDSD and presents the POST2 simulation flight dynamics support during the SFDT-1 launch, operations, and recovery.

NASA's Low Density Supersonic Decelerator (LDSD) project conducted its first Supersonic Flight Dynamics Test (SFDT-1) on June 28, 2014. Program to Optimize Simulated Trajectories II (POST2) was one of the flight dynamics codes used to simulate and predict the flight performance and Monte Carlo analysis was used to characterize the potential flight conditions experienced by the test vehicle. This paper compares the simulation predictions with the reconstructed trajectory of SFDT-1. Additionally, off-nominal conditions seen during flight are modeled in post-flight simulations to find the primary contributors that reconcile the simulation with flight data. The results of these analyses are beneficial for the pre-flight simulation and targeting of the follow-on SFDT flights currently scheduled for summer 2015.

NASA's Low Density Supersonic Decelerator (LDSD) program was established to identify, develop, and eventually qualify to Test [i.e. Technology] Readiness Level (TRL) - 6 aerodynamic decelerators for eventual use on Mars. Through comprehensive Mars application studies, two distinct Supersonic Inflatable Aerodynamic Decelerator (SIAD) designs were chosen that afforded the optimum balance of benefit, cost, and development risk. In addition, a Supersonic Disk Sail (SSDS) parachute design was chosen that satisfied the same criteria. The final phase of the multi-tiered qualification process involves Earth Supersonic Flight Dynamics Tests (SFDTs) within environmental conditions similar to those that would be experienced during a Mars Entry, Descent, and Landing (EDL) mission. The first of these flight tests (i.e. SFDT-1) was completed on June 28, 2014 with two more tests scheduled for the summer of 2015 and 2016, respectively. The basic flight design for all the SFDT flights is for the SFDT test vehicle to be ferried to a float altitude of 120 kilo-feet by a 34 thousand cubic feet (Mcf) heavy lift helium balloon. Once float altitude is reached, the test vehicle is released from the balloon, spun-up for stability, and accelerated to supersonic speeds using a Star48 solid rocket motor. After burnout of the Star48 motor the vehicle decelerates to pre-flight selected test conditions for the deployment of the SIAD system. After further deceleration with the SIAD deployed, the SSDS parachute is then deployed stressing the performance of the parachute in the wake of the SIAD augmented blunt body. The test vehicle/SIAD/parachute system then descends to splashdown in the Pacific Ocean for eventual recovery. This paper will discuss the development of both the test vehicle and the trajectory sequence including design trade-offs resulting from the interaction of both engineering efforts. In addition, the SFDT-1 nominal trajectory design and associated sensitivities will be discussed as well as an overview of the on-board flight software used to trigger and sequence the main flight events necessary to deploy the deceleration technologies. Finally, as-flown performance of the SFDT-1 system will be discussed.

The Low Density Supersonic Decelerator project's first Supersonic Flight Dynamics Test (SFDT) occurred on June 28, 2014, off the west coast of Kauai, Hawaii, over the Pacific Ocean. The test vehicle traveled to speeds above Mach 4 and to an altitude of over 200,000 feet. This flight, although classified as a test architecture shake-out flight, tested two technologies: a robotic class Supersonic Inflatable Aerodynamic Decelerator and a Supersonic Disksail Parachute. The reconstruction team was tasked with collecting all relevant pre-flight and flight data to accurately reconstruct the trajectory and technology performance during the science phase of the flight. Furthermore, the reconstruction team has been involved with reconstructing and exploring all aerodynamic and test vehicle properties that affected the entire flight phase. This reconstruction provided insight into the technology performance, which is a key deliverable for the LDSD project, as well as provided insight into lessons learned for subsequent SFDT flights, in the fields of data recovery, reconstruction, and pre-flight trajectory simulations.

The proposed Mars 2018 mission is a joint mission with the National Aeronautics and Space Administration (NASA) and the European Space Agency (ESA) that is currently in the early study phase. The objective would be to land a rover which caches a sample of Martian soil for retrieval and potential return to Earth on a subsequent mission. The Mars 2018 Entry, Descent, and Landing (EDL) system would largely leverage off of heritage from the Mars Science Laboratory (MSL) EDL system in an effort to minimize cost and schedule risk. Despite the desire to have high MSL heritage, a series of trade studies were performed to quantify the impact that incorporating EDL technologies would have on the landed mass. These new technologies include a trim tab, an increased diameter disk-gap-band (DGB) parachute over MSL, a large ringsail parachute, and a supersonic inflatable aerodynamic decelerator (SIAD).

In 2011, the Mars Science Laboratory (MSL) was launched in a mission to deliver the largest and most capable rover to date to the surface of Mars. A follow on MSL-derived mission, referred to as Mars 2018, is being proposed to launch in 2018. Mars 2018 is investigating performance enhancements of the Entry, Descent and Landing (EDL) system over that of its predecessor MSL mission of 2011. This paper will discuss the main elements of the proposed Mars 2018 EDL preliminary design that are being considered to increase performance on the entry phase of the mission. In particular, these elements are discussed with the goals of increasing the parachute deploy altitude to allow for more time margin during the subsequent descent and landing phases, increasing the entry mass, and reducing the delivery ellipse size at parachute deploy, through modifications in the entry reference trajectory design, vehicle's lift to drag ratio, parachute deploy trigger logic design, and the effect of additional navigation hardware.

Mars landers to date have flown ballistic entry trajectories with no trajectory control after the final maneuver before entry. 12Improvements in landing accuracies (from ~150 km from the target for Mars Pathfinder to ~30-40 km for MER and Phoenix) have been driven by approach navigation improvements. MSL will fly the first guided-entry trajectory to Mars, further improving accuracy to ~10-12 km from the target. For future missions, landing within ~100m is desired to assure landing safety close to a target of high scientific interest in irregular terrain, or to land near a previously landed asset. Improvements in approach navigation alone are not sufficient to achieve this requirement. If approach navigation error and IMU error are eliminated, the dominant error source is wind drift on the parachute, with map-tie error also significant. Correcting these errors requires terrain-relative navigation (TRN), which can be accomplished with passive imaging supplemented by radar for terrain sensing (with onboard navigation capable of processing measurements from IMU, imaging, and radar). Additionally, near-optimal-ΔV powered descent guidance is needed to minimize the amount of propellant required to reach the target. The capability to land within 100m can be applied in different landing modes depending on how much fuel is carried.

The Ares I-X Flight Test Vehicle was launched on October 28, 2009 and was the first and only test flight of NASA s two-stage Ares I launch vehicle design. The launch was successful and the flight test met all of its primary and secondary objectives. This paper discusses the stage separation and reentry trajectory analysis that was performed in support of the Ares I-X test flight. Pre-flight analyses were conducted to assess the risk of stage recontact during separation, to evaluate the first stage flight dynamics during reentry, and to define the range safety impact ellipses of both stages. The results of these pre-flight analyses were compared with available flight data. On-board video taken during flight showed that the flight test vehicle successfully separated without any recontact. Reconstructed trajectory data also showed that first stage flight dynamics were well characterized by pre-flight Monte Carlo results. In addition, comparisons with flight data indicated that the complex interference aerodynamic models employed in the reentry simulation were effective in capturing the flight dynamics during separation. Finally, the splash-down locations of both stages were well within predicted impact ellipses.

NASA senior management commissioned the Entry, Descent and Landing Systems Analysis Study in 2008 to identify and roadmap the Entry, Descent and Landing technology investments that the agency needed to make in order to successfully land large payloads at Mars for both robotic and human-scale missions. This paper summarizes the approach and top-level results from Year 2 of the Study, which focused on landing 1‐4 mt on Mars for robotic missions. Two separate studies were conducted in Year 2: the Mars Science Laboratory Improvement Study, which determined technology development program needs to support increases in landed payload and landed altitude beyond the Mars Science Laboratory capability using an Atlas V launch vehicle and the Exploration Feed-Forward Study, which examined a potential precursor mission using a Delta IV-H launch vehicle with landed payload in the 2‐4 mt range that would demonstrate key technologies needed for later human missions.

An algorithm for choosing the parachute deployment point has been developed which is designed to reduce delivery errors at powered descent ignition during Entry / Descent / Landing (EDL) on a body with an atmosphere (e.g. Mars), consequently reducing the propellant required to achieve precise landing at a preselected target. This algorithm is designed to improve on the previously developed “Smart Chute” deployment algorithm by modeling the lander’s trajectory during the parachute phase for improved targeting. Performance benefits are influenced by environmental factors, principally winds between chute deployment and ignition, which cause the lander to drift on the parachute.

Previous Mars landers have been able to land only within tens to hundreds of km of a target site. Principal sources of uncertainty are approach navigation, atmospheric modeling, and vehicle aerodynamics; additional (lesser) uncertainty sources are map-tie error and wind drift. The Mars Science Laboratory mission scheduled for 2009 launch will use guidance during hypersonic entry to improve this to ~10 km. To achieve "pinpoint landing" (within 100m) for future missions, ways of addressing the remaining error sources (approach navigation, wind drift and map-tie error) must be found. This work defines a "reference system design" for guidance, navigation, and control in future pinpoint landing missions, and assesses uncertainties and performance penalties associated with pinpoint landing using this reference system design