Jarret M. Lafleur

Extension of Traditional Entry, Descent, and Landing Technologies for Human Mars Exploration

The human exploration of Mars presents many challenges, not least of which is the task of entry, descent, and landing. Because human-class missions are expected to have landed payload masses on the order of 40 to 80 t, significant challenges arise beyond those of current robotic missions. This study uses parametric trade studies to provide insight into the feasibility of using Viking and Apollo heritage technologies to enable a human-class mission to Mars. The challenges encountered with human-class missions, as well as potential solutions, are highlighted through the results of parametric studies on vehicle size and mass. To populate the trade space, aerocapture, and entry-from-orbit analyses of 10 and 15-m diam aeroshells with a lift-to-drag ratio of 0.3 and 0.5 were investigated. The methodology developed to perform these trade studies represents a significant advancement in human Mars entry, descent, and landing system sizing. Numerous comparisons are made with past missions, both real and conceptual, and sources of discrepancies are discussed. Results indicate that in the limit, a crew capsule used only for entry from orbit could have an arrival mass as low as 20 t. For larger landed payloads, such as a 20-t surface power system, a vehicle with an arrival mass on the order of 80 t may be required. Finally, no feasible entry, descent, and landing systems were obtained with the capability to deliver more than approximately 25 t of landed payload to the Mars surface for arrivalmasses less than 100 t. This suggests that extension of traditional entry, descent, and landing technologies used for robotic exploration may be insufficient for human Mars exploration.

Exploring the F6 Fractionated Spacecraft Trade Space with GT-FAST

Released in July 2007, the Broad Agency Announcement for DARPA’s System F6 outlined goals for flight demonstration of an architecture in which the functionality of a traditional monolithic satellite is fulfilled with a fractionated cluster of free-flying, wirelessly interconnected modules. Given the large number of possible architectural options, two challenges facing systems analysis of F6 are (1) the ability to enumerate the many potential candidate fractionated architectures and (2) the ability to analyze and quantify the cost and benefits of each architecture. This paper applies the recently developed Georgia Tech F6 Architecture Synthesis Tool (GT-FAST) to the exploration of the System F6 trade space. Combinatorial analysis of the architectural trade space is presented, providing a theoretical contribution applicable to future analyses and clearly showing the explosion of the size of the trade space as the number of fractionatable components increases. Several output metrics of interest are defined, and Pareto fronts are used to visualize the trade space. The first set of these Pareto fronts allows direct visualization of one output against another, and the second set presents cost plotted against a Technique for Order Preference by Similarity to Ideal Solution (TOPSIS) score aggregating performance objectives. These techniques allow for the identification of a handful of Pareto-optimal designs from an original pool of over 3,000 potential designs. Conclusions are drawn on salient features of the resulting Pareto fronts, important competing objectives which have been captured, and the potential suitability of a particularly interesting design designated PF0248. A variety of potential avenues for future work are also identified.

GT-FAST: A Point Design Tool for Rapid Fractionated Spacecraft Sizing and Synthesis

In July 2007, DARPA issued a Broad Agency Announcement for the development of System F6, a flight demonstration of an architecture in which the functionality of a traditional monolithic satellite is fulfilled with a fractionated cluster of free-flying, wirelessly interconnected modules. Given the large number of possible architectural options, two challenges facing systems analysis of F6 are (1) the ability to enumerate the many potential candidate fractionated architectures and (2) the ability to analyze and quantify the cost and benefits of each architecture. One element necessary in enabling a probabilistic, valuecentric analysis of such fractionated architectures is a systematic method for sizing and costing the many candidate architectures that arise. The Georgia Tech F6 Architecture Synthesis Tool (GT-FAST) is a point design tool designed to fulfill this need by allowing rapid, automated sizing and synthesis of candidate F6 architectures. This paper presents the internal mechanics and some illustrative applications of GT-FAST. Discussed are the manner in which GT-FAST fractionated designs are specified, including discrete and continuous-variable inputs, as well as the methods, models, and assumptions used in estimating elements of mass, power, and cost. Finally, the paper concludes with sample outputs from GT-FAST for a notional fractionated architecture, an example of GT-FAST’s trade study capability, and a partial validation of GT-FAST against the Jason-2 and TIMED satellites. The ease with which GT-FAST can be adapted to new fractionated spacecraft applications is highlighted, and avenues for potential future expansion of GT-FAST are discussed.

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

*† ‡ ‡ ‡ ‡ , 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.

Probabilistic AHP and TOPSIS for multi-attribute decision-making under uncertainty

One challenging aspect in designing complex engineering systems is the task of making informed design decisions in the face of uncertainty.1,2 This paper presents a probabilistic methodology to facilitate such decision-making, in particular under uncertainty in decision-maker preferences. This methodology builds on the frequently-used multi-attribute decision-making techniques of the Analytic Hierarchy Process (AHP) and Technique for Order Preference by Similarity to Ideal Solution (TOPSIS), and it overcomes some typical limitations that exist in relying on these deterministic techniques. The methodology is divided into three segments, each of which consists of multiple steps. The first segment (steps 1–4) involves setting up the problem by defining objectives, priorities, uncertainties, design attributes, and candidate designs. The second segment (steps 5–8) involves applications of AHP and TOPSIS using AHP prioritization matrices generated from probability density functions. The third segment (steps 9–10) involves visualization of results to assist in selecting a final design. A key characteristic measured in these final steps is the consistency with which a design ranks among the top several alternatives. An example satellite orbit and launch vehicle selection problem illustrates the methodology throughout the paper.

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.

System-Level Feasibility Assessment of Microwave Power Beaming for Small Satellites

Although wireless power transmission to fulfill Earths energy needs has been a widely popularized application of microwave power beaming, one space application that remains relatively unexplored is power beaming between satellites. This paper provides a system-level analysis of the feasibility and limitations of microwave power beaming within a small-satellite cluster. This analysis consists of four parts using parametric models of spacecraft power as a function of 11 key design variables. In the first part, the existence of feasible designs is verified with a Monte Carlo design trade-space sweep. Next, a feasible baseline (reference) design is defined, and then sensitivity to individual variables is assessed. Finally, the design space is visualized with respect to six influential variables. Despite several optimistic assumptions, it is demonstrated that the small-satellite power-beaming design space is severely constrained. Feasible designs involve high transmission frequencies (greater than 33 GHz), large antenna diameters (greater than 0.93 m), and stringent proximity operations between satellites (within 740 m). Furthermore, full dependence of one small spacecraft on power provided by another is shown to be effectively infeasible. These results do suggest, however, that microwave power beaming may deserve consideration as an auxiliary or short-term emergency power mode for future small-satellite clusters.

Mars Entry Bank Profile Design for Terminal State Optimization

One challenge examined in NASA’s DRM 5.0 study is t hat of entry, descent, and landing (EDL) on Mars for high-ballistic-coefficient, human-class payloads. To define best-case entry scenarios for the evaluation of potential EDL system designs, a study is conducted to optimize the entry-to-terminal-state portion of EDL for a variety of entry velocities, vehicle ballistic coefficients (β), and lift-to-drag ratios (L/D). 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 and ten bank profile points to find maximum-final-altitude trajectories. A baseline set of optimizations is performed, as are fulllift-up and relaxed-deceleration-constraint sets fo r comparison. In total, an estimated 9 million trajectories are analyzed to yield 1800 opt imal trajectories. Parametric plots of maximum achievable altitude are shown, as are examp les of optimized trajectories. Characteristic vehicle contours are overlaid on the parametric plots, and conclusions are drawn on the feasibility of vehicles in the L/D vs. β design space. It is shown that entry bank angle control is highly deserving of consideration early in design, particularly for vehicles with midor high-L/D values, high entry velocities, and deceleration-li mited trajectories. Key conclusions are also drawn regarding trends in optimal bank profiles and in the constraints which impose particularly severe limits on the design of these trajectories.

Design of a long endurance Titan VTOL vehicle

Saturns moon Titan promises insight into numerous key scientific questions, many of which can be investigated only by in situ exploration of its surface and atmosphere. This paper presents research on a vertical takeoff and landing (VTOL) vehicle designed to conduct a scientific investigation of Titans atmosphere, clouds, haze, surface, and any possible oceans. Multiple options for vertical takeoff and horizontal mobility were considered. A helicopter was baselined because of its many advantages over other types of vehicles, particularly in that it has 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 mission operations, allowing it to sustain a long range, long duration mission that could travel over a thousand kilometers. The turbo expander power source can 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. This design is the first to investigate the implications of this potentially revolutionary technology on a Titan aerial vehicle

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.