No full-text available
To read the full-text of this research,
you can request a copy directly from the authors.
Precision Landing at Mars Using Discrete-Event Drag Modulation
Abstract
An entry, descent, and landing architecture capable of achieving Mars Science Laboratory-class landed accuracy (within 10 km of target) while delivering a Mars Exploration Rover-class payload to the surface of Mars is presented. The architecture consists of a Mars Exploration Rover-class aeroshell with a rigid, annular drag skirt. Maximum vehicle diameter, including drag skirt, is limited to be compatible with current launch-vehicle fairings. A single drag-skirt jettison event is used to control range during entry. Three-degree-of-freedom trajectory simulation is used in conjunction with Monte Carlo techniques to assess the flight performance of the proposed architecture. Results indicate that landed accuracy is competitive with preflight Mars Science Laboratory estimates, and peak heat rate and integrated heat load are significantly reduced relative to the Mars Exploration Rover entry system. Modeling parachute descent within the onboard guidance algorithm is found to remove range error bias present at touchdown; the addition of a range-based parachute deploy trigger is found to significantly improve landed accuracy.
... Practically, control of aerodynamic forces will improve the deceleration capability of the entry vehicle, which benefits fuel saving during the following powered descent phase. Existing entry guidance algorithms are traditionally tailored to the L/D ratio of the vehicle, including high L/D [8], [9], mid to low L/D [10], [11], and others [12], [13], [14]. A unified guidance method has been developed for a wide range of entry vehicles with varying lifting capabilities [15]. ...
... Moreover, the curve fitting results for P pressure and T temp with respect to the altitude h = Rh 0 − R m are written as T temp = −31 − 0.000998h, P pressure = 0.699 exp (−0.00009h) . Finding a polynomial curve to fit ρ with respect to h according to (14) is straightforward. The classical curve fitting method, such as the least square method, can find a polynomial function directly. ...
... The classical curve fitting method, such as the least square method, can find a polynomial function directly. However, using a polynomial to directly fit (14) will result in negative values at some points of the approximate air density function. Thus, (14) is firstly approximated by an exponential function, expressed as ...
This paper investigates the fuel-optimal guidance problem of the end-to-end human Mars entry, powered descent, and landing (EDL) mission. It applies a unified modeling scheme and develops a computationally efficient new optimization algorithm to solve the multi-phase optimal guidance problem. The end-to-end EDL guidance problem is first modeled as a multi-phase optimal control problem with different dynamics and constraints at each phase. Via polynomial approximation and discretization techniques, this multi-phase optimal control problem is then reformulated as a polynomial programming problem. By introducing intermediate variables and quadratic equality constraints, a polynomial program is equivalently converted into a nonconvex quadratically constrained quadratic program (QCQP). Then, a novel customized alternating direction method of multipliers (ADMM) is proposed to efficiently solve the large-scale QCQP with convergence proof to a local optimum under certain conditions on the algorithmic parameters. The fuel savings under the end-to-end human-Mars EDL guidance are verified by comparing to the fuel consumption using the separate phase guidance approach. Furthermore, the computational efficiency of the customized ADMM algorithm is validated by comparing to the state-of-the-art nonlinear programming method. The robustness of the customized ADMM algorithm is verified via extensive simulation cases with random initial conditions.
... Therefore, a numerical predictor-corrector approach, see [15][16][17] for example, that does not rely on a reference trajectory is a better choice for Mars entry with a SIAD. A predictor-corrector approach for Mars entry is investigated by Putnam and Braun [18] to control the drag-skirt jettison time which is also a discrete-event. In contrast to the SIAD, the drag-skirt is jettisoned during the entry to reduce the reference area. ...
... The lateral guidance is accomplished using previously developed bank reversal logic [24]. A range trigger, which has been found to be effective for the landing accuracy improvement [14,18,[25][26][27], is employed for the parachute deployment. The performance of the Mars entry guidance algorithm is assessed for various cases by numerical simulations. ...
... For example, NASA is currently investigating this technology further through the low-density supersonic decelerator program to land larger payload masses on the surface of Mars [4]. Another use of a SIAD is to provide discrete (or continuous) drag modulation as a means of landing site control [5][6][7][8]. In this case, SIAD deployment reduces the ballistic coefficient β, providing a means to directly control downrange. ...
... In this study, discrete-event drag modulation (SIAD deployment) is used to control vehicle downrange and the subsequent deployment conditions of the parafoil. For a more detailed discussion on drag modulation, refer to [8]. Both a fixed deployment altitude trigger and a predictor-corrector trigger were implemented in this study for range control evaluation. ...
The goal of this investigation is to understand the sizing and performance of supersonic inflatable aerodynamic decelerators for Earth-based sounding rocket applications. The recovery system under examination is composed of a supersonic inflatable aerodynamic decelerator and a guided parafoil system to achieve sub-100 m miss distances. Three supersonic inflatable aerodynamic decelerator configurations (tension cone, attached isotensoid, and trailing isotensoid) are examined using the metrics of decelerator mass, aerodynamic performance, and vehicle integration. In terms of aerodynamic performance, the tension cone is the preferred choice for the sizes investigated. The attached isotensoid was shown to be the most mass efficient decelerator, whereas the trailing isotensoid was found to be the more ideal decelerator for vehicle integration.Athree-degree-of-freedom trajectory simulation is used in conjunction with Monte Carlo uncertainty analysis to assess the landed accuracy capability of the proposed architectures. In 95% of the cases examined, the drag-modulated inflatable aerodynamic decelerator provides arrivals within the 10 km parafoil capability region, meeting the sub-100 m landed recovery goals. In 76% of the cases examined, the dragmodulated inflatable aerodynamic decelerator arrives within 5 km of this target zone. Copyright © 2014 by the American Institute of Aeronautics and Astronautics, Inc.
... Research has also been conducted in using drag modulation to achieve more precise landings and aerocaptures on Mars [3,15]. Drag modulation can also be used during the atmospheric reentry to reduce the aerodynamic loads [16] and to simplify the control during the descent [17]. ...
... The c 1 and c 2 terms in Eqs. (15) and (16) are constants resulting from the integration process of the indefinite integrals. A similar expression for RAAN Ω can be obtained by looking at the change in RAAN produced by the Earth's oblateness [23]: ...
The ability to control the location of a spacecraft atmospheric interface reentry has been traditionally accomplished using propulsion. A novel technique is presented here where a predefined point of atmospheric interface reentry is achieved by adjusting the aerodynamic drag of a spacecraft in a circular orbit. If this method is employed at a sufficiently high starting altitude, any ground-track point accessible by the orbit can be targeted. This method can be broken up into two different parts. The first consists of finding the decay profile that achieves the desired reentry location. The second consists of keeping the spacecraft within this nominal decay trajectory, given the atmospheric uncertainty and other perturbations. The two parts of this method are presented here in detail. Finally, a case study is provided to demonstrate how this method could work in a realistic scenario and to evaluate its performance. The case study shows that a reentry point can be targeted with a 3σ error of less than 200 km, using a typical Global Positioning System for navigation. Finally, a discussion of potential applications is provided. Copyright © 2014 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.
... For example, NASA is currently investigating this technology further through the low-density supersonic decelerator program to land larger payload masses on the surface of Mars [4]. Another use of a SIAD is to provide discrete (or continuous) drag modulation as a means of landing site control [5][6][7][8]. In this case, SIAD deployment reduces the ballistic coefficient β, providing a means to directly control downrange. ...
... In this study, discrete-event drag modulation (SIAD deployment) is used to control vehicle downrange and the subsequent deployment conditions of the parafoil. For a more detailed discussion on drag modulation, refer to [8]. Both a fixed deployment altitude trigger and a predictor-corrector trigger were implemented in this study for range control evaluation. ...
This paper presents an assessment of a supersonic inflatable aerodynamic decelerator for use on a sounding rocket payload bus structure for a high-altitude sample return mission. Three decelerator configurations, the tension cone, attached isotensoid, and the trailing isotensoid, were examined on the metrics of decelerator mass, aerodynamic performance, and vehicle integration. The mass calculated for diameters ranging from approximately 0.5 to 1.1 m with a baseline dimater 0.9 m is shown to be similar between each configuration. However, the attached isotensoid configuration is shown to be the least mass solution. Greater than 50% drag performance degradation results when the attachment point recessed from the forebody of the bus structure. Using multiattribute decision making techniques, the trailing isotensoid, is identified to be the most advantageous decelerator optionfor use in this application.
... In practice, however, the requirement that the drag skirt deploy in subsonic conditions severely limits the total achievable control authority, to the extent that this approach has little or no merit for this application. This is because by the time the vehicle reaches subsonic speeds it has already dissipated almost all of its energy and is at a low altitude (about 10 km in this case), Mars [47,48]. However, the drag skirt for the current SHIELD concept would not structurally or thermally withstand such conditions ‡ . ...
Motivated by a need for lower-cost planetary science missions to Mars, this study considers the problem of co-delivering a network of small rough landers to the Martian surface such that the probes are placed on different entry trajectories by a single carrier spacecraft without requiring translational maneuvers between probe deployments. The Small High Impact Energy Landing Device is used as a reference design, and a flight-mechanics analysis is performed to ensure requirements are met under the influence of relevant uncertainties. A linearized targeting method is developed and applied to design probe jettison velocities for a regional probe network. Monte Carlo analysis shows that a regional network approximately 100 km in scale could be passively co-delivered with limited deformation of the network shape despite the influence of relevant uncertainties, and linearized targeting errors are quantified. Nonlinear numerical optimization is applied and enables the design of probe jettisons for co-delivery of larger-scale networks. Additional Monte Carlo analyses quantify the rate at which delivery error increases with network scale.
... Drag-modulation control may be possible during aerobraking atmospheric passes by using already-articulated solar panels as controllable drag areas. Drag modulation has been studied for aerocapture and entry, where drag areas are typically jettisoned in discrete, one-time events [4,10]. However, other studies have considered the use of continuously-variable geometry to modulate the heat, drag, or acceleration loads of a blunted body in an entry trajectory [11]. ...
An energy depletion guidance algorithm for shallow, high-altitude atmospheric flight, consistent with aerobraking atmospheric passes, is proposed. The guidance algorithm is based on optimal control solutions to the energy minimization problem and uses variable drag area for trajectory control achieved through articulated solar panels. The algorithm includes three control modes to limit heat rate, heat load, or both while attempting to maximize energy depletion. A closed-form approximate solution of the equations of motion for aerobraking atmospheric passes is used to generate steering commands in real time. Numerical simulation of aerobraking at Mars illustrates the performance of the algorithm modes for a range of initial conditions. Results indicate that the guidance algorithm increases nominal energy depletion by approximately 1600% relative to the uncontrolled case over a single aerobraking pass, equivalent to an increase in ΔV of approximately 1600%. Monte Carlo simulation analyses show that dispersed flight performance is consistent with nominal performance; the standard deviation of the depleted energy is at maximum 13.8% of the mean, and the maximum difference between the mean and the maximal energy depletion solution is lower than 1.4%.
... The strategy has been studied for direct entry at Mars, providing MSL-class landed accuracy on the order of 10 km. 12 Drag-modulation in aerocapture provides an effective theoretical corridor width (TCW) analogous to lifting trajectories based on the ratio of ballistic coefficients before and after the discrete drag-modulation event. 13,14 Less obvious systems-level benefits of drag-modulation include passively stable aerodynamic configurations and mitigation of launch-vehicle payload fairing restrictions on the vehicle geometry. ...
... Such vehicles would require a large ratio for the designed low and high values of ballistic coefficients. Drag modulations can be used for both entry and aerocapture at Mars [14,15]. ...
... Despite this simplicity, previous analyses have shown that discrete drag modulation systems are capable of achieving accuracy competitive with that of lift-modulation systems. 1,4,5 Additionally, drag-modulation systems typically have low ballistic coefficients, resulting in a more benign aerothermal environment than that experienced by lifting aerocapture concepts, enabling drag modulation systems to utilize less expensive and lighter weight thermal protection system materials. ...
Small satellites may provide a low-cost platform for targeted science investigations in the Mars system. With current technology, small satellites require ride shares with larger orbiters to capture into orbit, limiting the range of orbits available to small satellite mission designers. Successful development of a small satellite aerocapture capability would allow small satellite mission designers to choose the orbit most appropriate for a science investigation while enabling small satellite ride shares on any mission to Mars. A generic small satellite aerocapture system is assessed for use at Mars across a range of small satellite payloads, approach trajectories, and destinations in the Mars system. The aerocapture system uses drag modulation for trajectory control to ensure successful orbit insertion. Analyses include assessment of the sensitivity of the entry corridor size to the ballistic-coefficient ratio, the effectiveness of real-time aerocapture guidance and control algorithms, aerocapture system-level impacts of different target orbits, and development of requirements and recommendations for the development of a small satellite aerocapture system. Results indicate that a discrete drag-modulation aerocapture system may provide an orbit-insertion capability for small satellites with modest propulsion requirements.
... The remaining 8 employed some combination of aeroshell or IAD at earlier phases in the flight when such methods are advantageous. Discrete drag modulation events, for example, solve some of the targeting problems caused by aerodynamic methods by jettisoning the decelerating element once the vehicle is determined to be on target [52] [53]. The sudden increase in gives the vehicle resilience to atmospheric perturbations, but means the transition to high-flight is determined by uncontrollable day-of-flight factors. ...
Supersonic retropropulsion (SRP) is the use of retrorockets to decelerate during
atmospheric flight while the vehicle is still traveling in the supersonic/hypersonic
flight regime. In the context of Mars exploration, subsonic retropropulsion has a
robust flight heritage for terminal landing guidance and control, but all supersonic
deceleration has, to date, been performed by non-propulsive (i.e. purely aerodynamic)
methods, such as aeroshells and parachutes.
Extending the use of retropropulsion from the subsonic to the supersonic regime
has been identified as an enabling technology for high mass humans-to-Mars architectures.
However, supersonic retropropulsion still poses significant design and control
challenges, stemming mainly from the complex interactions between the hypersonic
engine plumes, the oncoming air flow, and the vehicle’s exterior surface. These interactions
lead to flow fields that are difficult to model and produce counter intuitive
behaviors that are not present in purely propulsive or purely aerodynamic flight.
This study will provide an overview of the work done in the design of SRP systems.
Optimal throttle laws for certain trajectories will be derived that leverage
aero/propulsive e↵ects to decrease propellant requirements and increase total useful
landing mass. A study of the mass savings will be made for a 10 mT reference vehicle
based on a propulsive version of the Orion capsule, followed by the 100 mT ellipsoid
vehicle assumed by NASA’s Mars Design Reference Architecture.
Several critical technologies are needed for a human mission to Mars that require considerable further development. These include reliable environmental control and life support systems (ECLSS), mitigation of radiation and low gravity effects, large-scale entry, descent and landing, utilization of indigenous planetary resources, and human factors associated with long durations in confined space. The reliability of ECLSS systems falls far short of requirements. NASA has made progress in understanding radiation effects but as more information accrues, the problem appears worse. Human factors appear to be a major problem, hardly investigated to date. A vital need for a human mission to Mars is aero-assisted entry, descent, and precision landing (EDL) of massive payloads. There is no experience base for landing payloads with mass of multi-tens of mT. Modeling by the Georgia Tech team indicates that the mass of EDL systems will be considerably greater than that assumed by NASA Design Reference Missions. Recent studies have not clarified the picture. Developing, testing, and validating such massive entry systems will require a two-decade program with a significant investment.
Entry flight performance is assessed for single-stage discrete-event drag-modulation trajectory control on Mars with two different real-time guidance algorithms: a heuristic velocity trigger and numerical predictor–corrector. Three degree-of-freedom simulation and Monte Carlo techniques are used to determine flight performance across a range of mission and vehicle design parameters. Trends are identified in flight performance across different initial conditions, ballistic-coefficient ratios, and target ranges. Results indicate that both guidance algorithms can provide landing accuracy better than 10 km across likely vehicle properties and entry conditions. The heuristic velocity trigger performs nearly as well as the numerical predictor–corrector for low-ballistic-coefficient ratios and entry flight-path angles steeper than approximately –18 deg. The numerical predictor–corrector provides consistent flight performance across all feasible mission and vehicle parameters with accuracy better than approximately 5 km. Overall, single-stage discrete-event drag-modulation trajectory control is a feasible option for accurate landings for ballistic coefficients below approximately 140 kg/m2 and landed altitudes above 0 km.
This paper presents an analysis of the hypersonic separation dynamics for a planetary entry vehicle and a deployable drag area that is jettisoned during planetary entry. The first portion of the paper identifies separation times for a range of vehicle parameters and flight conditions which includes vehicle size, drag area size, and jettison velocity. For a given entry trajectory, minimum separation time corresponds to a jettison event at maximum dynamic pressure. Results show that vehicles with a larger initial ballistic coefficients require less time to achieve a separation distance of one drag area aft radius, in addition to a smaller range of possible separation times. Analysis is performed in the second portion of this paper to determine if recontact occurs between the two bodies after jettison. Results indicate recontact is most likely to occur at jettison conditions with low dynamic pressures, large angles of attack, and angle of attack rates, and with small differences in ballistic coefficient between the entry vehicle and drag area. Sensitivity analysis is performed to determine which conditions are most important to the success of the jettison event. Results show that the difference in ballistic coefficient between entry vehicle and drag area is the most influential parameter in determining the recontact-free jettison envelope. This envelope can be manipulated through moving the drag area center of gravity off of the axis of symmetry. Overall, results indicate a successful jettison, with no recontact, is possible and likely for typical blunt body spherecone entry trajectories and attitude dynamics. The chance of recontact can be minimized by jettisoning the drag area closer to maximum dynamic pressure.
Small satellites may provide a low-cost platform for targeted science investigations in the Mars system. With current technology, small satellites require ride shares with larger orbiters to capture into orbit, limiting the range of orbits available to small satellite mission designers. Successful development of an independent orbit insertion capability for small satellites, using aerocapture, would allow small satellite mission designers to choose the orbit most appropriate for a science investigation while enabling small satellite ride shares on any mission to Mars. A generic small satellite drag-modulation aerocapture system is assessed for use at Mars across a range of approach trajectories and destinations in the Mars system. Analyses include assessment of the sensitivity of the entry corridor size over different atmospheric conditions, a comparison of velocity-trigger and numerical predictor-corrector guidance schemes for drag modulation, and aerocapture flight performance assessment via Monte Carlo techniques. A special focus is placed on four baseline missions: a low-altitude Mars mapping orbit, Phobos and Deimos flyby/rendezvous, and areosynchronous orbit. Results indicate that aerocapture may decrease the orbit insertion system mass fraction by 30% or more with respect to fully propulsive options.
Discrete-event drag-modulation trajectory control is assessed for planetary entry using the closed-form Allen-Eggers solution to the equations of motion. A control authority metric for drag-modulation trajectory control systems is derived. Closed-form analytical relationships are developed to assess range divert capability and to identify jettison condition constraints for limiting peak acceleration and peak heat rate. Closed-form relationships are also developed for drag-modulation systems with an arbitrary number of stages.
A number of critical technologies are needed for a human mission to Mars that require considerable further development. These include life support: environmental control and life support systems (ECLSS), mitigation of radiation and low gravity effects, providing abort options, potential utilization of indigenous planetary resources, and human factors associated with long durations in confined space. While significant progress was made on ECLSS prior to 2005, there is little indication of progress in the past decade. NASA has made progress in understanding radiation effects but as more information accrues, the problem appears worse. Work on artificial gravity seems moribund. Use of simulated habitats in remote areas on Earth is helping to gradually understand issues associated with confined space. A vital need for a human mission to Mars is aero-assisted entry, descent and landing (EDL) of massive payloads. There is no experience base for landing payloads with mass of multi-tens of mT. Modeling by the Georgia Tech team indicates that the mass of EDL systems will be considerably greater than that assumed by NASA Design Reference Missions. Nevertheless, aero assisted EDL requires far less mass than EDL based on propulsion, and use of propulsion for EDL is probably unaffordable. Developing, testing and validating such massive entry systems will require a two-decade program with a significant investment. Based on past performance, NASA does not appear to have the discipline to follow through on such a program.
This study proposes an on-line predictor-corrector reentry guidance algorithm that satisfies path and no-fly zone constraints for hypersonic vehicles with a high lift-to-drag ratio. The proposed guidance algorithm can generate a feasible trajectory at each guidance cycle during the entry flight. In the longitudinal profile, numerical predictor-corrector approaches are used to predict the flight capability from current flight states to expected terminal states and to generate an on-line reference drag acceleration profile. The path constraints on heat rate, aerodynamic load, and dynamic pressure are implemented as a part of the predictor-corrector algorithm. A tracking control law is then designed to track the reference drag acceleration profile. In the lateral profile, a novel guidance algorithm is presented. The velocity azimuth angle error threshold and artificial potential field method are used to reduce heading error and to avoid the no-fly zone. Simulated results for nominal and dispersed cases show that the proposed guidance algorithm not only can avoid the no-fly zone but can also steer a typical entry vehicle along a feasible 3D trajectory that satisfies both terminal and path constraints.
Landing heavier payloads with pinpoint precision on the surface of Marsis aknown challenge for future missions to Mars. Supersonic retropropulsion has been proposed as a means to deliver higher-mass payloads to the surface, traditionally with the sole intent of deceleration. By adding range control capability during the supersonic retropropulsion maneuverby meansofthrust vectoring, it has been found thatasubstantial amountofpropellant can be saved in a precision landing scenario when compared with traditional architectures. Decreasing the propellant mass necessary for the mission increases the amount of payload mass that can be brought to the surface. Propellant mass savings greater than 30% are possible if thrust vectoring is unconstrained during the supersonic phase of flight. Propellant mass fraction is found to be sensitive to the divert direction and also the altitude and flight-path angle at ignition, favoring low altitudes and shallow flight-path angles. Decreased nozzle cant angles and aerodynamic drag preservation have also been found to reduce propellant usage. Copyright © 2015 by A. B. Mandalia and R. D. Braun. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission.
Future robotic and human missions to Mars require improved landed precision and increased payload mass. Two architectures that seek to meet these requirements using supersonic propulsive diverts are proposed in this paper: one utilizing a high-altitude propulsive divert and another with thrust vectoring during supersonic retropropulsion. Low ballistic coefficient entry vehicles decelerate high in the thin Mars atmosphere and may be used to deliver higher-mass payloads to the surface. A high-altitude supersonic propulsive divert maneuver is proposed as a means of precision landing for low ballistic coefficient entry vehicles that decelerate to supersonic speeds at altitudes of 20-60 km. This divert maneuver compares favorably to state-of-the-art precision landing architectures with range accuracy on the order of 100 m while saving over 30% in propellant mass. Architectures which utilize hypersonic vehicles with ballistic coefficients of 10 kg/m2 can possibly land within 500 m of a target with this maneuver alone. Supersonic retropropulsion has also been proposed as a means to deliver higher-mass payloads to the surface, and thrust vectoring during supersonic retropropulsion can save a substantial amount of fuel in a precision landing scenario. Propellant mass savings greater than 30% are possible if thrust vectoring is unconstrained during the supersonic phase of flight.
The closed-form analytical solution to the equations of motion for ballistic entry developed by Allen and Eggers is presented using modern nomenclature and is extended and enhanced through several new developments. A method of choosing an appropriate constant flight-path angle is identified.Analytical limits are proposed that bound the domain of applicabilityof the approximation.Closed-formexpressions for range andtime are developedthat are consistentwiththe assumptions in theAllen-Eggers approximation. Collectively, the improvements address key weaknesses in the original approximate solution.Assessment of the accuracy of the approximate solution relative to the planar equations ofmotion shows that the extended and enhanced Allen-Eggers solution provides good accuracy across a wide range of ballistic coefficients at Earth with initial flight-path angles steeper than about-7 deg. In some instances, the expression developed for range-to-go may be accurate enough for use in onboard real-time guidance and targeting systems.
In this investigation, a parametric study for the preliminary design of an Earth atmospheric dust collection and recovery mission has been conducted. The scientific goal of this mission is to sample and recover mesospheric dust and particulate matter. Suborbital flight trajectories, vehicle configurations, and deceleration technologies were analyzed using conceptual models. The trajectory is shown to be driven by the science objective (sample collection at 45 km to 85 km in altitude) and the target dust and particulate matter size. Preliminary vehicle configuration results indicate an insensitivty to landing dispersion and show a spacecraft-dependent relation to total heating. From the initial results, the design space is pruned and three reference mission architectures are defined-one which utilizes a standard disk-gap-band parachute and two that utilize supersonic inflatable aerodynamic decelerators. With use of the inflatable aerodynamic decelerator, drag modulation is shown to be able to reduce the landed uncertainty in downrange by approximately 6.8 km at the 95% confidence level.
The Mars Science Laboratory (MSL) is a NASA rover mission that will be launched in late 2011 and will land on Mars in August of 2012. This paper describes the analyses performed to validate the navigation system for launch, interplanetary cruise, and approach. MSL will use guidance during its descent into Mars in order to minimize landing dispersions, and therefore will be able to use smaller landing zones that are closer to terrain of high scientific interest. This will require a more accurate delivery of the spacecraft to the atmospheric entry interface, and a late update of the state of the spacecraft at entry. During cruise and approach the spacecraft may perform up to six trajectory correction maneuvers (TCMs), to target to the desired landing site with the required flight path angle at entry. Approach orbit determination covariance analyses have been performed to evaluate the accuracy that can be achieved in delivering the spacecraft to the entry interface point, and to determine how accurately the state of the spacecraft can be predicted to initialize the guidance algorithm. In addition, a sensitivity analysis has been performed to evaluate which factors most contribute to the improvement or degradation of the navigation performance, for both entry flight path angle delivery and entry state knowledge.
Trailing Ballute Aerocapture offers the potential to obtain orbit insertion around a planetary body at a fraction of the mass of traditional methods. This allows for lower costs for launch, faster flight times and additional mass available for science payloads. The technique involves an inflated ballute (balloon-parachute) that provides aerodynamic drag area for use in the atmosphere of a planetary body to provide for orbit insertion in a relatively benign heating environment. To account for atmospheric, navigation and other uncertainties, the ballute is oversized and detached once the desired velocity change (Delta V) has been achieved. Analysis and trades have been performed for the purpose of assessing the feasibility of the technique including aerophysics, material assessments, inflation system and deployment sequence and dynamics, configuration trades, ballute separation and trajectory analysis. Outlined is the technology development required for advancing the technique to a level that would allow it to be viable for use in space exploration missions.
The Mars Exploration Rover mission will be the next opportunity for surface exploration of Mars in January 2004. Two rovers will be delivered to the surface of Mars using the same entry, descent, and landing scenario that was developed and successfully implemented by Mars Pathfinder. This investigation describes the trajectory analysis that was performed for the hypersonic portion of the MER entry. In this analysis, a six-degree-of-freedom trajectory simulation of the entry is performed to determine the entry characteristics of the capsules. In addition, a Monte Carlo analysis is also performed to statistically assess the robustness of the entry design to off-nominal conditions to assure that all entry requirements are satisfied. The results show that the attitude at peak heating and parachute deployment are well within entry limits. In addition, the parachute deployment dynamics pressure and Mach number are also well within the design requirements.
An overview of several important aerodynamics challenges new to the Mars Science Laboratory (MSL) entry vehicle are presented. The MSL entry capsule is a 70° sphere-cone based on the original Mars Viking entry capsule. Due to payload and landing accuracy requirements, MSL will be flying at the highest lift-to-drag ratio of any capsule sent to Mars (L/D = 0.24). The capsule will also be flying a guided entry, performing bank maneuvers, a first for Mars entry. The system's mechanical design and increased performance requirements require an expansion of the MSL flight envelope beyond those of historical missions. In certain areas, the experience gained by Viking and other recent Mars missions can no longer be claimed as heritage information. New analysis and testing is required to ensure the safe flight of the MSL entry vehicle. The challenge topics include: hypersonic gas chemistry and laminar-versus-turbulent flow effects on trim angle, a general risk assessment of flying at greater angles-of-attack than Viking, quantifying the aerodynamic interactions induced by a new reaction control system and a risk assessment of recontact of a series of masses jettisoned prior to parachute deploy. An overview of the analysis and tests being conducted to understand and reduce risk in each of these areas is presented. The need for proper modeling and implementation of uncertainties for use in trajectory simulation has resulted in a revision of prior models and additional analysis for the MSL entry vehicle. The six degree-of-freedom uncertainty model and new analysis to quantify roll torque dispersions are presented.
The twin Mars Exploration Rover missions landed successfully on Mars' surface in January of 2004. Both missions used a parachute system to slow the rover's descent rate from supersonic to subsonic speeds. Shortly after parachute deployment, the heat shield, which protected the rover during the hypersonic entry phase of the mission, was jettisoned using push-off springs. Mission designers were concerned about the heat shield recontacting the lander after separation, so a separation analysis was conducted to quantify risks. This analysis was used to choose a proper heat shield ballast mass to ensure successful separation with a low probability of recontact. This paper presents the details of such an analysis, its assumptions, and the results. During both landings, the radar fortuitously locked onto the heat shield and measured its distance, as it descended away from the lander. This data is presented and is used to validate the heat shield separation/recontact analysis.
A software tool for the prediction of the aero -thermodynamic environments of conceptual aerospace configurations is presented. The vehic le geometry is defined using unstructured, triangulated surface meshes. For subsonic Mach numbers a fast, unstructured, multi -pole panel code is coupled with a streamline tracing formulation to define the viscous surface solution. For supersonic and hypers onic Mach numbers, various independent panel methods are coupled with the streamline tracing formulation, an attachment line detection method , and stagnation -attachment line heating models to define the viscous aero -thermal environment.
An end-to-end simulation of the Mars Science Laboratory (MSL) entry, descent, and landing (EDL) sequence was created at the NASA Langley Research Center using the Program to Optimize Simulated Trajectories II (POST2). This simulation is capable of providing numerous MSL system and flight software responses, including Monte Carlo-derived statistics of these responses. The MSL POST2 simulation includes models of EDL system elements, including those related to the parachute system. Among these there are models for the parachute geometry, mass properties, deployment, inflation, opening force, area oscillations, aerodynamic coefficients, apparent mass, interaction with the main landing engines, and offloading. These models were kept as simple as possible, considering the overall objectives of the simulation. The main purpose of this paper is to describe these parachute system models to the extent necessary to understand how they work and some of their limitations. A list of lessons learned during the development of the models and simulation is provided. Future improvements to the parachute system models are proposed.
Aerocapture using a towed, inflatable ballute system has been shown to provide signifi-cant performance advantages compared to traditional technologies, including lower heating rates and accommodation of larger navigational uncertainties. This paper extends previous results by designing a ballute aerocapture separation algorithm that can operate in a more realistic Titan atmospheric model based on TitanGRAM. This model incorporates both latitudinal variability as well as noisiness in the density profile.
Drag-modulation flight control may provide a simple method for controlling energy during aerocapture. Several drag-modulation flight-control system options are discussed and evaluated, including single-stage jettison, two-stage jettison, and continuously variable drag-modulation systems. Performance is assessed using numeric simulation with real-time guidance and targeting algorithms. Monte Carlo simulation is used to evaluate system robustness to expected day-of-flight uncertainties. Results indicate that drag-modulation flight control is an attractive option for aerocapture systems at Mars, where low peak heat rates enable the use of lightweight inflatable drag areas. Aerocapture using drag modulation at Titan is found to require large drag areas to limit peak heat rates to nonablative thermal-protection system limits or advanced lightweight ablators. The large gravity well and high peak heat rates experienced during aerocapture at Venus make drag-modulation flight control unattractive when combined with a nonablative thermal-protection system. Significantly larger drag areas or advances in fabric-based material thermal properties are required to improve feasibility at Venus.
Aerothermal analyses of the Phoenix entry vehicle in the Martian Atmosphere are presented. Two Navier-Stokes flow solvers are used to predict the environments experienced by the vehicle along several design trajectories. The majority of the solutions are computed as axisymmetric in order to yield higher, and therefore more conservative, heating rates. Several three-dimensional cases are also computed to assess the heating enhancement caused by surface singularities and the complex aftbody vortices. When the thermal protection system was assessed in a separate analysis in response to the computed aeroheating environments, the prebuilt design is found perform successfully. Copyright © 2008 by the American Institute of Aeronautics and Astronautics, Inc.
The Mars Pathfinder spacecraft will enter the Martian atmosphere directly from the interplanetary trajectory, with a velocity of up to 7.35 km/s. The definition of the nominal entry trajectory and the accurate determination of potential trajectory dispersions are necessary for the design of the Pathfinder entry, descent, and landing system. Monte Carlo numerical simulations have been developed to quantify the range of possible entry trajectories and attitude profiles. The entry trajectory requirements and constraints are discussed, and the design approach and uncertainties used in the Monte Carlo analysis are described. Three-degree-of-freedom and six-degree-of-freedom trajectory results are compared. The Monte Carlo analysis shows that the Mars Pathfinder parachute will be deployed within the required ranges of dynamic pressure, Mach number, and altitude, over a 3 sigma range of trajectories. The Pathfinder 3 sigma landing ellipse is shown to he roughly 50 x 300 km.
This study presents a means of explicit guidance for ballistic entry using an improved method of matched asymptotic expansions. The trajectory of ballistic entry into a planetary atmosphere is still an important and often critical phase of a mission. In the paper, feedback control via drag modulation is used to guide the vehicle during the atmospheric entry, whereas a matched asymptotic solution for the entry trajectory is available to aim the target. The feedback control ensures the stability of a trajectory around the nominal trajectory by compensating for the non-linear terms in the motion of the vehicle. Using the improved method of matched asymptotic expansions, the control algorithms for the guidance law are derived explicitly and tested against the 1976 U.S. Standard Atmosphere. Simulation results indicate that the control algorithms can effectively control the trajectories in the lower atmosphere under the targeting dispersions of atmospheric variations.
Two primary simulations have been developed and are being updated for the Mars Science Laboratory entry, descent, and landing. The high-fidelity engineering end-to-end entry, descent, and landing simulation is based on NASA Langley Research Center's Program to Optimize Simulated Trajectories II and the end-to-end real-time, hardwaee-in-the-loop simulation test bed, which is based on NASA Jet Propulsion Laboratory's Dynamics Simulator for Entry, Descent, and Surface landing. The status of these Mars Science Laboratory entry, descent, and landing end-to-end simulations at this time is presented. Various models, capabilities, as well as validation and verification for these simulations, are discussed.
The Mars Pathfinder mission provides the next opportunity for scientific exploration of the surface of Mars following a 7.6 km/s direct entry: In support of this effort, a six-degree-of-freedom trajectory analysis and aerodynamic characteristic assessment are performed to demonstrate vehicle flyability and to quantify the effect that each of numerous uncertainties has upon the nominal mission profile. The entry vehicle is shown to be aerodynamically stable over a large portion of its atmospheric flight. Two low angle-of-attack static instabilities (freestream velocities of about 6.5 and 3.5 km/s) and a low angle-of-attack dynamic instability (supersonic) are identified and shown to cause bounded increases in vehicle attitude. The effects of center-of-gravity placement, entry attitude, vehicle roll rate, aerodynamic misprediction, and atmospheric uncertainty on the vehicle attitude profile and parachute deployment conditions are quantified. A Monte Carlo analysis is performed to statistically assess the combined impact of multiple off-nominal conditions on the nominal flight characteristics. These results suggest that there is a 99.7% probability that the peak attitude throughout the entry will be less than 8.5 deg, the peak heating attitude mill be below 6.2 deg, and the attitude at parachute deployment will be less than 3.9 deg.
This paper gives compact formulae for the direct and inverse solutions of geodesics of any length. Existing formulae have been recast for efficient programming to conserve space and reduce execution time. The main feature of the new formulae is the use of nested equations for elliptic terms. Both solutions are iterative.
The United States has successfully landed five robotic systems on the surface of Mars. These systems all had landed mass below 0.6 metric tons (t), had landed footprints on the order of hundreds of km and landed at sites below -1 km MOLA elevation due the need to perform entry, descent and landing operations in an environment with sufficient atmospheric density. Current plans for human exploration of Mars call for the landing of 40-80 t surface elements at scientifically interesting locations within close proximity (10's of m) of pre-positioned robotic assets. This paper summarizes past successful entry, descent and landing systems and approaches being developed by the robotic Mars exploration program to increased landed performance (mass, accuracy and surface elevation). In addition, the entry, descent and landing sequence for a human exploration system will be reviewed, highlighting the technology and systems advances required
In 2012, during the Entry, Descent, and Landing (EDL) of the Mars Science Laboratory (MSL) entry vehicle, a 21.5 m Viking-heritage, Disk-Gap-Band, supersonic parachute will be deployed at approximately Mach 2. The baseline algorithm for commanding this parachute deployment is a navigated planet-relative velocity trigger. This paper compares the performance of an alternative range-to-go trigger (sometimes referred to as “Smart Chute”), which can significantly reduce the landing footprint size. Numerical Monte Carlo results, predicted by the POST2 MSL POST End-to-End EDL simulation, are corroborated and explained by applying propagation of uncertainty methods to develop an analytic estimate for the standard deviation of Mach number. A negative correlation is shown to exist between the standard deviations of wind velocity and the planet-relative velocity at parachute deploy, which mitigates the Mach number rise in the case of the range trigger.
The method developed in NASA TN D-319 for studying the atmosphere entry of vehicles with varying aerodynamic forces has been applied to obtain a closed-form solution for the motion, heating, range, and variation of the vehicle parameter m/C(D)A for nonlifting entries during which the rate of increase of deceleration is limited. The solution is applicable to vehicles of arbitrary weight, size, and shape, and to arbitrary atmospheres. Results have been obtained for entries into the earth's atmosphere at escape velocity during which the maximum deceleration and the rate at which deceleration increases were limited. A comparison of these results with those of NASA TN D-319, in which only the maximum deceleration was limited, indicates that for a given corridor depth, limiting the rate of increase of deceleration and the maximum deceleration requires an increase in the magnitude of the change in M/C(D)A and results in increases in maximum heating rate, total heat absorbed at the stagnation point, and range.
The stagnation-point convective heat transfer to an axisymmetric blunt body for arbitrary gases in chemical equilibrium was investigated. The gases considered were base gases of nitrogen, oxygen, hydrogen, helium, neon, argon, carbon dioxide, ammonia, and methane and 22 gas mixtures composed of the base gases. Enthalpies ranged from 2.3 to 116.2 MJ/kg, pressures ranged from 0.001 to 100 atmospheres, and the wall temperatures were 300 and 1111 K. A general equation for the stagnation-point convective heat transfer in base gases and gas mixtures was derived and is a function of the mass fraction, the molecular weight, and a transport parameter of the base gases. The relation compares well with present boundary-layer computer results and with other analytical and experimental results. In addition, the analysis verified that the convective heat transfer in gas mixtures can be determined from a summation relation involving the heat transfer coefficients of the base gases. The basic technique developed for the prediction of stagnation-point convective heating to an axisymmetric blunt body could be applied to other heat transfer problems.
This paper presents the complete analysis of the problem of minimum-fuel aeroassisted transfer between coplanar elliptical orbits in the case where the orientation of the final orbit is free for selection in the optimization process. The comparison between the optimal pure propulsive transfer and the idealized aeroassisted transfer, by several passages through the atmosphere, is made. In the case where aeroassisted transfer provides fuel saving, a practical scheme for its realization by one passage is proposed. The maneuver consists of three phases: a deorbit phase for nonzero entry angle, followed by an atmospheric fly-through with variable drag control and completed by a postatmospheric phase. An explicit guidance formula for drag control is derived and it is shown that the required exit speed for ascent to the final orbit can be obtained with a very high degree of accuracy.
Stagnation-point radiative heating rate expressions are presented for use in air and an approximate Martian atmosphere consisting of 97 percent CO2 and 3 percent N2. Thermochemical equilibrium is assumed throughout. The flight conditions and body dimensions that are modeled are representative of both manned and unmanned missions to Mars and return to earth. Comparisons between the heating rates computed using the expressions presented here and independent computations yielded maximum differences of about 20 to 30 percent.
The author has analyzed the use of a light-weight inflatable hypersonic drag device, called a ballute, for flight in planetary atmospheres, for entry, aerocapture, and aerobraking. Studies to date include Mars, Venus, Earth, Saturn, Titan, Neptune and Pluto, and data on a Pluto lander and a Mars orbiter will be presented to illustrate the concept. The main advantage of using a ballute is that aero, deceleration and heating in atmospheric entry occurs at much smaller atmospheric density with a ballute than without it. For example, if a ballute has a diameter 10 times as large as the spacecraft, for unchanged total mass, entry speed and entry angle,the atmospheric density at peak convective heating is reduced by a factor of 100, reducing the heating by a factor of 10 for the spacecraft and a factor of 30 for the ballute. Consequently the entry payload (lander, orbiter, etc) is subject to much less heating, requires a much reduced thermal. protection system (possibly only an MLI blanket), and the spacecraft design is therefore relatively unchanged from its vacuum counterpart. The heat flux on the ballute is small enough to be radiated at temperatures below 800 K or so. Also, the heating may be reduced further because the ballute enters at a more shallow angle, even allowing for the increased delivery angle error. Added advantages are less mass ratio of entry system to total entry mass, and freedom from the low-density and transonic instability problems that conventional rigid entry bodies suffer, since the vehicle attitude is determined by the ballute, usually released at continuum conditions (hypersonic for an orbiter, and subsonic for a lander). Also, for a lander the range from entry to touchdown is less, offering a smaller footprint. The ballute derives an entry corridor for aerocapture by entering on a path that would lead to landing, and releasing the ballute adaptively, responding to measured deceleration, at a speed computed to achieve the desired orbiter exit conditions. For a lander an accurate landing point could be achieved by providing the lander with a small gliding capacity, using the large potential energy available from being subsonic at high altitude. Alternatively the ballute can be retained to act as a parachute or soft-landing device, or to float the payload as a buoyant aerobot. As expected, the ballute has smaller size for relatively small entry speeds, such as for Mars and Titan, or for the extensive atmosphere of a low-gravity planet such as Pluto. Details of a ballute to place a small Mars orbiter and a small Pluto lander will be given to illustrate the concept. The author will discuss presently available ballute materials and a development program of aerodynamic tests and materials that would be required for ballutes to achieve their full potential.
The Mars Surveyor 2001 project will send an orbiter, a lander, and a rover to Mars in the 2001 opportunity. The lander will demonstrate precision landing at Mars by utilizing aggressive approach navigation and hypersonic aeromaneuvering. The guided entry will result in a landed footprint that is an order of magnitude smaller than the Mars Pathfinder and Mars Polar Lander ballistic entry footprints. This paper will focus on the interplanetary navigation strategy that will decrease entry errors and reduce the size of the landed footprint.
Drag Modulation and Celestial Mechanics
- P H Rose
- J E Hayes
Rose, P. H., and Hayes, J. E., " Drag Modulation and Celestial Mechanics, " Advances in the Astronautical Sciences, Vol. 8, Plenum Press, New York, 1963.
The Next Generation of Mars-GRAM and its Role in the Autonomous Aerobraking Develop-ment Plan
- H L Justh
- C G Justus
- H S Ramey
Justh, H. L., Justus, C. G., and Ramey, H. S., " The Next Generation of Mars-GRAM and its Role in the Autonomous Aerobraking Develop-ment Plan, " AAS/AIAA Astrodynamics Specialist Conference, Paper 11-478, 2011.
Mars Exploration Rover Heat Shield Recontact Analysis, " 21st AIAA Aerodynamic Decelerator Systems Technology Conference and Seminar, AIAA Paper
- B Raiszadeh
- P N Desai
- R Michelltriee
Raiszadeh, B., Desai, P. N., and Michelltriee, R., " Mars Exploration Rover Heat Shield Recontact Analysis, " 21st AIAA Aerodynamic Decelerator Systems Technology Conference and Seminar, AIAA Paper 2011-2584, May 2011. C. Kluever Associate Editor 138 PUTNAM AND BRAUN Downloaded by UNIVERSITY OF CALIFORNIA -DAVIS on February 5, 2015 | http://arc.aiaa.org | DOI: 10.2514/1.A32633
Entry, Descent and Landing Systems Analysis Study
- A M D Cianciolo
- J L Davis
- W C Engelund
- D R Komar
- E M Queen
- J A Samareh
- D W Way
- T A Zang
- J G Murch
- S A Krizan
- A D Olds
- R W Powell
- J D Shidner
- D Kinney
- M K Mcguire
- J O Arnold
- M A Covington
- R R Sostaric
- C H Zumwalt
- E G Llama
Cianciolo, A. M. D., Davis, J. L., Engelund, W. C., Komar, D. R., Queen, E. M., Samareh, J. A., Way, D. W., Zang, T. A., Murch, J. G., Krizan, S. A., Olds, A. D., Powell, R. W., Shidner, J. D., Kinney, D., McGuire, M. K., Arnold, J. O., Covington, M. A., Sostaric, R. R., Zumwalt, C. H., and Llama, E. G., " Entry, Descent and Landing Systems Analysis Study, " NASA TM-2011-217055, Feb. 2011.