Trading Robustness Requirements in Mars Entry Trajectory Design

One of the most important metrics characterizing an atmospheric entry trajectory in preliminary design is the size of its predicted landing ellipse. Often, requirements for this ellipse are set early in design and significantly influence both the expected scientific return from a particular mission and the cost of development. Requirements typically specify a certain probability level (6-level) for the prescribed ellipse, and frequently this latter requirement is taken at 36. However, searches for the justification of 36 as a robustness requirement suggest it is an empirical rule of thumb borrowed from non-aerospace fields. This paper presents an investigation into the sensitivity of trajectory performance to varying robustness (6-level) requirements. The treatment of robustness as a distinct objective is discussed, and an analysis framework is presented involving the manipulation of design variables to effect trades between performance and robustness objectives. The scenario for which this method is illustrated is the ballistic entry of an MSL-class Mars entry vehicle. Here, the design variable is entry flight path angle, and objectives are parachute deploy altitude performance and error ellipse robustness. Resulting plots show the sensitivities between these objectives and trends in the entry flight path angles required to design to these objectives. Relevance to the trajectory designer is discussed, as are potential steps for further development and use of this type of analysis

Angle of Attack Modulation for Mars Entry Terminal State Optimization

From the perspective of atmospheric entry, descent, and landing (EDL), one of the most foreboding destinations in the solar system is Mars due in part to its exceedingly thin atmosphere. To benchmark best possible scenarios for evaluation of potential Mars EDL system designs, a study is conducted to optimize the entry-to-terminal-state portion of EDL for a variety of entry velocities and vehicle masses, focusing on the identification of potential benefits of enabling angle of attack modulation. The terminal state is envisioned as one appropriate for the initiation of terminal descent via parachute or other means. A particle swarm optimizer varies entry flight path angle, ten bank profile points, and ten angle of attack profile points to find maximum-final-altitude trajectories for a 10 30 m ellipsled at 180 different combinations of values for entry mass, entry velocity, terminal Mach number, and minimum allowable altitude. Parametric plots of maximum achievable altitude are shown, as are examples of optimized trajectories. It is shown that appreciable terminal state altitude gains (2.5-4.0 km) over pure bank angle control may be possible if angle of attack modulation is enabled for Mars entry vehicles. Gains of this magnitude could prove to be enabling for missions requiring high-altitude landing sites. Conclusions are also drawn regarding trends in the bank and angle of attack profiles that produce the optimal trajectories in this study, and directions for future work are identified

StarRunner: A Single-Stage-to-Orbit, Airbreathing, Hypersonic Propulsion System

40th AIAA/ASME/SAE/ASEE Joint Propulsion Conference And Exhibit Fort Lauderdale, FL, July 11-14, 2004.In response to the request for proposal (RFP) for the 2003 AIAA Undergraduate Team Engine Design Competition, the FAS Propulsion Design team from the Georgia Institute of Technology presents StarRunner: A Single-Stage-to-Orbit (SSTO), Airbreathing, Hypersonic Propulsion System. Low-cost, highly reliable access to low-Earth orbit (LEO) and the International Space Station (ISS) is an area of continuing research and debate. StarRunner is proposed to supplement a notional Crew Transfer Vehicle through the ability to deliver a 25,000 lb payload to the ISS. The horizontal takeoff/horizontal landing (HTHL) vehicle makes use of a turbine-based combined cycle (TBCC) propulsion system consisting of 14 low-bypass-ratio turbofan engines and a dual-mode ramjet/scramjet propulsion system for high-speed flight. The vehicle also takes advantage of ultra-high-temperature ceramic thermal protection materials and uses hydrogen fuel for regenerative cooling of engine components. StarRunner is compatible with standard runways, with a gross takeoff weight of approximately 1,000,000 lbs, and has a cost per pound to orbit of approximately $825/lb. This advanced, fully reusable space transport vehicle and integrated propulsion system design demonstrates student efforts to understand issues facing the space launch community. Future enabling and enhancing technologies for TBCC SSTO launch vehicles are explored and analyzed. The final StarRunner design addresses and proposes several innovative solutions to traditional problems

A Systematic Concept Exploration Methodology Applied to Venus In Situ Explorer

This presentation was part of the session : Probe Missions to the Giant Planets, Titan and VenusSixth International Planetary Probe WorkshopA critical task in any design process is the initial conversion of customer or program objectives into a baseline system architecture. This task becomes particularly important for space exploration systems that have unique requirements which, in many cases, have never been met before. A useful tool to the space systems engineer would be a methodology which helps to make this objectives-to-design conversion more systematic and efficient. Presented in this paper is such a methodology frequently used at the Georgia Institute of Technology, and in this paper the methodology is applied to initial concept formulation for the Venus In Situ Explorer (VISE) mission. VISE is one of six New-Frontiers-class missions to occur within the next 30 years that NASA addresses in its Solar System Exploration Roadmap. VISE is envisioned as an aerial mission that will study Venus' atmospheric composition as well as descend briefly to the surface to acquire samples for later analysis at more benign altitudes. Common to both VISE and its successor, Venus Mobile Explorer, is the challenge to operate under the extreme temperatures (about 730 K) and pressures (about 90 atm) present at the Venusian surface. In order to establish a baseline mission and vehicle concept for VISE, the methodology presented here begins with problem definition and the generation of functional and operational architectures. Customer requirements and engineering targets are set through an established set of tools known as the seven management and planning tools and through the use of a quality function deployment (QFD). A morphological matrix is used to identify 12.4 billion potential solutions in the concept space. From this concept space, six representative designs are chosen to demonstrate how alternatives from the morphological matrix may be ranked through multi-attribute decision making (MADM) techniques such as the Technique for Order Preference by Similarity to Ideal Solution (TOPSIS) and Pugh concept selection matrices. Two of the six concepts are eliminated based on these MADM techniques, and the remaining four concepts are recognized as requiring more in-depth study to allow definitive rankings to be assigned. A notional modeling and simulation framework for this problem is formed which could be used to complete such an in-depth, quantitative study. This paper principally serves to illustrate an example of how a systematic objectives definition, concept generation, and downselection methodology can be applied to advanced interplanetary missions (specifically in the example of Venus In Situ Explorer). The methodology and tools presented here are shown as a helpful guide and addition to the toolbox of the space systems engineer during the advanced planning stages of design.Georgia Institute of Technology ; National Institute of Aerospac

Daedalon: A Revolutionary Morphing Spacecraft Design for Planetary Exploration

1st AIAA Space Exploration Conference January 2005, Orlando FL.The product of a study sponsored by the NASA Institute for Advanced Concepts (NIAC), Daedalon is a spacecraft design baselined for Mars which utilizes morphing wing technology to achieve the design objective of a standard, flexible architecture for unmanned planetary exploration. This design encompasses a detailed vehicle mass and power sizing study for the Daedalon lander as well as its cruise stage and entry backshell. A cost estimation and comparison study is also performed, and qualitative functionality comparisons are made between Daedalon and other Mars lander and airplane designs. Quantitative comparisons of gross mass and range are also made, including the results of scaling an existing Mars aerial vehicle design to match Daedalon functionality. Altogether, the Daedalon launch mass is found to be 896 kg for a 12 kg payload capacity. If five such vehicles are produced, it is found that the per-mission cost can be as low as $224 million. Given the necessary morphing wing technology development, it is concluded that the Daedalon design may be a feasible and cost-effective approach to planetary exploration 20-40 years in the future

Entry Descent and Landing Challenges of Human Mars Exploration

29th AAS Guidance and Control Conference February 2006, Breckenridge, CO.Near-term capabilities for robotic spacecraft include a target of landing 1 - 2 metric ton payloads with a precision of about 10 kilometers, at moderate altitude landing sites (as high as +2 km MOLA). While challenging, these capabilities are modest in comparison to the requirements for landing human crews on Mars. Human Mars exploration studies imply the capability to safely land 40 - 80 metric ton payloads with a precision of tens of meters, possibly at even higher altitudes. New entry, descent and landing challenges imposed by the large mass requirements of human Mars exploration include: (1) the potential need for aerocapture prior to entry, descent and landing and associated thermal protection strategies, (2) large aeroshell diameter requirements, (3) severe mass fraction restrictions, (4) rapid transition from the hypersonic entry mode to a descent and landing configuration, (5) the need for supersonic propulsion initiation, and (6) increased system reliability. This investigation explores the potential of extending robotic entry, descent and landing architectures to human missions and highlights the challenges of landing large payloads on the surface of Mars

Sizing of an Entry, Descent, and Landing System for Human Mars Exploration

AIAA Space 2006 Conference September 2006, San Jose, CA.The human exploration of Mars presents many challenges, not least of which is the task of entry, descent, and landing (EDL). Because human-class missions are expected to have landed masses on the order of 40 to 80 metric tons, significant challenges arise that have not been seen to date in robotic missions. This study provides insight into the challenges encountered as well as potential solutions through parametric trade studies on vehicle size and mass. Aerocapture and entry-from-orbit analyses of 10 and 15 m diameter aeroshells with a lift-to-drag ratio of 0.3 or 0.5 were investigated. Results indicate that in the limit, a crew capsule used only for descent could have an initial mass as low as 20 t. For larger landed payloads, such as a 20 t surface power system, a vehicle with an initial mass on the order of 80 t may be required. In addition, no feasible EDL systems were obtained with the capability to deliver more than approximately 25 t of landed payload to the Mars surface for initial masses less than 100 t. This suggests that an aeroshell diameter of 15 m may not be sufficient for human Mars exploration

Design of a Long Endurance Titan VTOL Vehicle

IEEE Aerospace Conference March 2006, Big Sky, MT.Saturn's moon Titan promises insight into many key scientific questions, many of which can be investigated only by in situ exploration of the surface and atmosphere of the moon. This research presents a vertical takeoff and landing (VTOL) vehicle designed to conduct a scientific investigation of Titan's atmosphere, clouds, haze, surface, and any possible oceans. In this investigation, multiple options for vertical takeoff and horizontal mobility were considered. A helicopter was baselined because of its many advantages over other types of vehicles, namely access to hazardous terrain and the ability to perform low speed aerial surveys. Using a nuclear power source and the atmosphere of Titan, a turbo expander cycle produces the 1.9 kW required by the vehicle for flight and operations, allowing it to sustain a long range, long duration mission that could traverse the majority of Titan. Such a power source could increase the lifespan and quality of science for planetary aerial flight to an extent that the limiting factor for the mission life is not available power but the life of the mechanical parts. Therefore, the mission could potentially last for years. This design is the first to investigate the implications of this potentially revolutionary technology on a Titan aerial vehicle