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Eight fixed-wing reusable horizontal landing booster point design concepts are presented and compared on the basis of weight, cost, technical difficulty, and availability date. The eight vehicle types considered are all basically two-stage systems with a lifting body reusable second stage, with all vehicles normalized to place 40,000 lbs. payload in orbit. All flight vehicles are fully recoverable and capable of flying back and landing at the launch site. Vehicle types discussed are vertical take-off horizontal landing rockets, sled launched horizontal take-off rockets, runway launched horizontal take-off rockets, air breathing first stages, combined air breathing and rocket first stages, oxidizer collection concepts, supersonic combustion ramjets, and in-flight refueling vehicles. Each of these vehicle types is depicted in the paper and its design and performance characteristics are discussed.

The paper compares the size, weight and performance characteristics of hydraulic, pneumatic and electro-mechanical actuators for general aviation flight controls. Mathematical models for each type are presented. Actuator designs for specific control tasks in realistic light airplane applications are compared with each other. It is shown that electro-mechanical actuators utilizing the recently developed samarium-cobalt technology have significant advantages in terms of size and weight requirements while yielding very fast response.

The INTELSAT VIII / VIIIA program consists of six communication satellites (of two different designs) being produced by Lockheed Martin Astro Space for the International Telecommunications Satellite Organization (INTELSAT). Two spacecraft level thermal vacuum test facilities were required to test the first four spacecraft due to schedule constraints. The two facilities selected were the East Windsor, NJ (EW) 35 foot chamber and the 39 foot chamber at Lockheed Martin's Valley Forge, PA (VF) facility. INTELSAT 801 was tested in the EW 35 foot chamber which utilizes eight independently controlled shrouds to perform IR (hot wall) testing over an environmental temperature range of -180 to +80°C. INTELSAT 802 was tested in the VF 39 foot chamber which consists of shrouds flooded with LN2 at -180°C. Cal rods were used to independently control the thermal environments of the six spacecraft faces for this facility.

In recent years, a tremendous interest has spawned towards adapting Lithium-Ion battery technology for aircraft applications. Lithium-Ion technology is already being used in some military aircraft (e.g., the F-22, F-35 and the B-2) and it has also been selected as original equipment for large commercial aircraft (e.g., the Airbus A380 and Boeing B787). The advantages of Lithium-Ion technology over Lead-Acid and Nickel-Cadmium technologies are higher specific energy (Wh/kg) and energy density (Wh/L), and longer cycle life. Saving weight is especially important in aircraft applications, because it can boost fuel economy and increase mission capability. Disadvantages of Lithium-Ion technology include higher initial cost, limited calendar/float life, inferior low temperature performance, and more severe safety hazards. This paper will present a direct comparison of a 24-Volt, 28Ah Lead-Acid and a 24-volt, 28Ah Lithium-Ion aircraft battery.

A study of indirect lightning effects on carbon composite structures with internal tubing is performed using the ‘ab initio’ Maxwell equation codes MAFIA and Microwave Studio (MWS). The modeling is performed both in the time and frequency domain by a finite difference method that can accommodate anisotropic media. Both time and frequency domains are used to better reflect the actual testing performed on aircraft. Solutions in frequency domain also help to overcome limitations of the time domain calculations. Time calculations cannot be performed for pulses longer than 100–200 μs due to the Courant condition and computer time limitations and hence, low frequency resonances of the system could be missed. Three dimensional frequency domain calculations are available with MAFIA and MWS, while time domain results are available in MAFIA, MWS and EMA3D.

Multiphase CFD simulations of air and water play a critical role in aircraft icing analysis. Specifically for air data sensors mounted near the front of an aircraft, simulations that predict the concentration of water surrounding an aircraft fuselage are necessary for understanding their performance in icing conditions. Those simulations can aid in sensor design and placement, and are central for defining critical conditions to test during icing qualification campaigns. There are several methods available in CFD that solve a multiphase flow field. Two of the most common methods used are Lagrangian and Eulerian. While these methods are similar, important differences can be viewed in the results, specifically in how the water shadow zones are predicted. This paper compares a Lagrangian and Eulerian CFD method for solving a multiphase flow field, and assesses their performance for use for analyzing installation locations and critical icing conditions of air data probes.

A comparison of various numerical tools and techniques was performed for calculating the lightning indirect effects to composite structures and internal systems. This paper is a summary of the initial comparison results. Detailed results of each technique considered are given in additional separate papers presented during this conference. The modeling considered current distributions over and within composite surfaces and the coupling of current and voltages to internal systems such as wire bundle cables and hydraulic and fuel tubes. The models were compared to each other and to measured data from low level swept continuous wave (LLCW) tests performed on two test fixtures. Other features of the codes such as run time, ease of use, computer requirements, availability of documentation and technical support, etc. are compared as well.

Several quite different waste-water processing subsystems have been tested for space mission applications. Mission choice, and choice of priority for further development, involves processor requirements for (electrical) power, launch and resupply weights, and space occupied, and developmental factors of reliability and reparability. For each subsystem one can identify theoretical power, etc., requirements, and current realistic requirements, and potential realistic subsystem requirements. Compared are reverse osmosis, vapor-compression-distillation, and TIMES processes, in terms of power, weight, and space requirements. Discussions of electrodialysis and CELSS are included as components of water treatment systems.

The space exploration enterprise that will lead to human exploration on Mars requires pressure suit system capabilities and characteristics that change significantly over time and between different missions and mission phases. These capabilities must be provided within tight budget constraints and severely limited launch mass and volume, and at a pace that supports NASA's over-all exploration timeline. As a result, it has not been obvious whether the use of a single pressure suit system (like Apollo) or combinations of multiple pressure suit designs (like Shuttle) will offer the best balance among life cycle cost, risk, and performance. Because the answer to this question is pivotal for the effective development of pressure suit system technologies that will met NASA's needs, ILC and Hamilton Sundstrand engineers have collaborated in an independent study to identify and evaluate the alternatives.

Three lift/propulsion concepts for a V/STOL supersonic combat aircraft have been compared. The intention was to show the effect of the propulsion system on aircraft weight and size, performance, and life cycle costs for: 1 Vectored thrust with Plenum Chamber Burning (bypass air augmentation) 2 Lift engines and a lift/cruise reheated turbofan 3 A reheated lift/cruise turbofan with a remote augmented lift system (RALS) For a postulated deck-launched intercept mission, the vectored thrust propulsion system with Plenum Chamber Burning gives the smallest and cheapest aircraft having the required performance. In addition, for a given take-off ground run the vectored thrust powered aircraft has the longest fighter escort mission radius.

The purpose of this study was to evaluate and identify the most effective techniques for self-guided remote training. Nine subjects were trained on the use of a personal videoconferencing system and the use of a Web-based multimedia program. Each subject performed two tasks with both training techniques: one assembly task and one instrumentation task. It was found that task times using the multimedia application were greater than the task times using videoconferencing. However, for complex tasks, participants preferred multimedia training. With multimedia, participants preferred to use video clips over diagrams and photos during assembly tasks. However, no preference was found during instrumentation tasks.

Wheat plants were grown at water potentials in the root zone of -0.4, -3.0, and -5.0 kPa in root modules with various porous membranes through which the nutrient solution was delivered. Root modules contained plants grown during 49 days on different types of porous membranes: ceramic porous tubes with diameters of 10 mm or 22 mm, a porous titanium plate, in a compartment with a porous ceramic tube in perlite and in a 2.5 cm layer of perlite which covered a porous titanium plate. Root modules containing perlite showed much higher dry mass plants in yield than plants in root modules without perlite. A drop in water potential resulted in growth inhibition in all of the modules, especially in the tests without perlite. Design characteristics of the modules significantly affected the root distribution volume. These results may provide additional information in the design of root modules for future space plant growth chambers.

Cooperation between Russia and the United States on manned spaceflight has led to unprecedented openness, resulting in the ability to now compare the characteristics of environmental control/life support hardware selected to generate oxygen (O2) by water electrolysis for space station applications. This comparison in this paper focuses on the characteristics that have the greatest effect on the cost of assembling and maintaining the hardware in space: launch weight, volume, power consumption, resupply requirements and maintenance labor.

Flight temperatures were compared to thermal balance test correlation data for two spacecraft. Results show that locations that were correlated with ground test data did not necessarily have similar correlation with flight data. Thermal models that were correlated extremely well showed the tendency toward better flight correlation. Thermal model correlation was successful in eliminating large temperature discrepancies in flight and providing a necessary understanding of the thermal design. Furthermore, flight correlation data verified the importance of the continued use of an ±11°C thermal uncertainty margin in spacecraft thermal design.

This paper presents the results of two static tests of structures in use on certificated aircraft and a description of finite element models of these structures. Comparisons of the results of the static tests and the predictions of the finite element analyses are given. Problems encountered in the development of the finite element models are presented, and solutions to these problems are discussed. Implementation, applications, and functional aspects of finite element analysis are also discussed. This paper does not cover the basic mathematical development of finite analysis, nor dynamic analysis applications.

The objective of this investigation was to measure and document the existence of any significant differences in physical performance under operational conditions between the Launch Entry Suit (LES) and the new Advanced Crew Escape Suit (ACES). The LES is a partial pressure suit currently worn by astronauts during the launch and entry phases of Shuttle missions. The ACES is a full pressure suit under consideration as a replacement for the LES. One prototype ACES has been fabricated and was used in this investigation. This report presents the results of three tests conducted with six subjects to allow a comparative evaluation of the two suits. The three tests included a reach envelope test, a strength test, and a treadmill test. The reach envelope test measured and compared the maximum hand displacements during horizontal and vertical reaches of both left and right arms in one-g conditions.

West Virginia University's Mechanical and Aerospace Engineering Department is studying the benefits of continuous payload volume in transforming projectiles. Continuous payload volume is the single largest vacancy in a vehicle that may be utilized. Currently there is a market for transforming projectiles, which are gun launched (or tube launched) vehicles stowed in an initial configuration; which deploy wings once exiting the launcher to become small unmanned aircraft. WVU's proposed design uses a helical hinge, which allows the wing sections to be externally stowed outside the UAV's fuselage. Additionally, the design positions the vehicles wing sections sub-bore (or smaller than the guns internal diameter), and flush (smooth and planer) to the surface of the fuselage. The typical transforming winged projectile design considered, stores its wing sections along the center axis of the fuselage. This bisects the payload space and limits the continuous payload carrying potential.

This paper investigates the use of an industry standard tube and closure method with the aim of replacing the existing glass tube. In order to accredit the tube the long-term stability of the samples was investigated in terms of accuracy, precision, and robustness with actual and control samples in the submarine environment and with control samples in the laboratory.

The Orion Crew and Service Modules are often compared to the Apollo Command and Service Modules due to their similarity in basic mission objective: both were dedicated to getting a crew to lunar orbit and safely returning them to Earth. Both spacecraft rely on the environmental control, life support and active thermal control systems (ECLS/ATCS) for the basic functions of providing and maintaining a breathable atmosphere, supplying adequate amount of potable water and maintaining the crew and avionics equipment within certified thermal limits. This assessment will evaluate the driving requirements for both programs and highlight similarities and differences. Further, a short comparison of the two system architectures will be examined including a side by side assessment of some selected system's hardware mass.