Multidisciplinary Design Techniques Applied to Conceptual Aerospace Vehicle Design

Multidisciplinary design optimization (MDO) is an emerging discipline within aerospace engineering. Its goal is to bring structure and efficiency to the complex design process associated with advanced aerospace launch vehicles. Aerospace vehicles generally require input from a variety of traditional aerospace disciplines - aerodynamics, structures, performance, etc. As such, traditional optimization methods cannot always be applied. Several multidisciplinary techniques and methods were proposed as potentially applicable to this class of design problem. Among the candidate options are calculus-based (or gradient-based) optimization schemes and parametric schemes based on design of experiments theory. A brief overview of several applicable multidisciplinary design optimization methods is included. Methods from the calculus-based class and the parametric class are reviewed, but the research application reported focuses on methods from the parametric class. A vehicle of current interest was chosen as a test application for this research. The rocket-based combined-cycle (RBCC) single-stage-to-orbit (SSTO) launch vehicle combines elements of rocket and airbreathing propulsion in an attempt to produce an attractive option for launching medium sized payloads into low earth orbit. The RBCC SSTO presents a particularly difficult problem for traditional one-variable-at-a-time optimization methods because of the lack of an adequate experience base and the highly coupled nature of the design variables. MDO, however, with it's structured approach to design, is well suited to this problem. The result of the application of Taguchi methods, central composite designs, and response surface methods to the design optimization of the RBCC SSTO are presented. Attention is given to the aspect of Taguchi methods that attempts to locate a 'robust' design - that is, a design that is least sensitive to uncontrollable influences on the design. Near-optimum minimum dry weight solutions are determined for the vehicle. A summary and evaluation of the various parametric MDO methods employed in the research are included. Recommendations for additional research are provided

Flight Performance of the Inflatable Reentry Vehicle Experiment 3

The Inflatable Reentry Vehicle Experiment 3 (IRVE-3) launched July 23, 2012, from NASA Wallops Flight Facility (WFF) on a Black Brant XI suborbital sounding rocket and successfully performed its mission, demonstrating the survivability of a hypersonic inflatable aerodynamic decelerator (HIAD) in the reentry heating environment and also illustrating the effect of an offset center of gravity on the HIAD's lift-to-drag ratio. IRVE-3 was a follow-on to 2009's IRVE-II mission, which demonstrated exo-atmospheric inflation, reentry survivability - without significant heating - and the aerodynamic stability of a HIAD down to subsonic flight conditions. NASA Langley Research Center is leading the development of HIAD technology for use on future interplanetary and Earth reentry missions

State-of-the-art clinical assessment of hand function

We have assembled a multi-disciplinary team of engineers, surgeons, clinicians and neuroscientists from Johns Hopkins School of Medicine and Western University to develop a new device for assessing hand function. It will be capable of sensitively measuring fingertip forces across all five fingers and along all movement directions. Then we can use this device to develop and validate a clinical hand assessment for patients with brain injuries.https://ir.lib.uwo.ca/brainscanprojectsummaries/1005/thumbnail.jp

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

Costing & economics for advanced launch vehicles

Issued as NASA grant project report, and Final report, Project no. E-16-N9

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

Modeling Approach for Analysis and Optimization of a Long-Duration Mars Airplane

AIAA MA&O Conference Albany, NY, May 30-June 1, 2004.The goal of this study was to determine the best system level modeling tool for the design of a long endurance Mars airplane mission, and to use this tool to determine the best configuration for the aircraft. The mission model was built in the design framework ModelCenter. User-driven fixed point iteration (FPI), optimizer based decomposition (OBD) and a hybrid method were implemented. Convergence difficulties were discovered in the OBD and hybrid methods. The user-driven FPI method produced the most reliable results, but required the most time. A combination of the hybrid and user-driven FPI methods were used to perform a technology study in which five different propulsion systems were examined: a bipropellant rocket, a battery powered propeller, a direct methanol fuel cell powered propeller, and beamed solar and microwave powered propeller systems. The direct methanol fuel cell proved to be the best onboard power system for a long endurance airplane and the solar beamed power system showed potential for indefinite flight