David Myre

Aerodynamics of an Aerofoil in Transonic Ground Effect: Numerical Study at Full-scale Reynolds Numbers

The potential positive effects of ground proximity on the aerodynamic performance of a \Ving or aerofoil have long been established, but at transonic speeds the fonnation of shock waves betvveen the body and the ground plane would have significant consequences. A numerical study of the aerodynamics of an lu\E2822 aerofoil section in ground eiiect flight was conducted at freestream Mach numbers from 0·5 to 0·9, at a range of ground clearances and angles of incidence. It \Vas found that in general the aerofoils lifting capability was still improved with decreasing ground clearance up until the point at >,vhich a lower surface shock wave formed (most commonly at the loviest clearances). The critical Mach number for the section \Vas reached considerably earlier in ground effect than it \vould be in freest ream, and the buffet boundary vvas therefore also reached at an earlier stage. The tlowfields observed were relatively sensitive to changes in any given variable, and the lower surface shock had a destabilising effect on the pitching characteristics ofthe Viing, indicating that sudden changes in both altii11de and attitude would be experienced during sustained transonic Hight close to the ground plane. Since ground proximity hastens the lower surface shock formation, no gain in aerodynamic efficiency can be

Exhaust Gas Analysis of a Vortex Oxidizer Injection Hybrid Rocket Motor

Previously a vortex oxidizer injection hybrid rocket motor was developed at the U.S. Naval Academy to investigate possible enhancements to fuel regression rate due to injection method. The motor was developed to operate with a conventional axial injector and a novel vortex injector design using high density polyethylene for fuel and a gaseous oxygen oxidizer. The motor was instrumented with pressure transducers and high-speed gas analyzers to measure exhaust gas carbon dioxide, carbon monoxide, oxygen and unburned hydrocarbons. After multiple tests of both injector configurations measurements for the vortex case indicated fuel regression was more evenly distributed across the fuel grain except at the head end, the average regression rate and chamber pressure were higher and exhaust gas analysis indicated a lower oxidizer-fuel ratio. In order to characterize the rocket chamber and compare with the gas measurements, a rocket chamber simulation was developed. The simulation is a one dimensional Lagrangian combustion analysis developed using Matlab and flame mechanisms available in Cantera, an open source gas/chemical analysis package which included an applicable combustion mechanism for ethylene. The model solves a series coupled ordinary differential equations, which include species balance, energy conservation and conservation of mass. The outputs include species mass fraction, temperature, velocity, density and specific gas constant. The system of equations solved quickly and provided realistic results. Because of low velocities in the chamber, the full kinetics model and equilibrium model differed only slightly. Also comparison with measured carbon dioxide was not consistent with the experimentally determined oxidizer-fuel ratios and requires further study. In this iteration of the model heat transfer were neglected.

In-flight performance of III-V multi-junction solar cells from the Forward Technology Solar Cell Experiment

The Materials on the International Space Station Experiments (MISSE) present a unique opportunity in space science by offering a low-cost platform to expose materials directly to the space environment on the International Space Station (ISS). MISSE experiments consist of a “suitcase” like package known as the “Passive Experiment Carrier” (PEC) that can be carried by astronauts and mounted externally to the ISS. The 5th MISSE payload (MISSE-5) contained both passive and active experiments. The Forward Technology Solar Cell Experiment (FTSCE) on MISSE-5 measured current-voltage (I–V) characteristics on 36 solar cells of various types. Over 1500 I–V curves were recorded on each cell during a 13-month period. This paper analyses the results for all the III–V multi-junction cells flown, including state-of-the-art space qualified cells and next generation metamorphic cells.

Science of opportunity: Heliophysics on the FASTSAT mission and STP-S26

The FASTSAT spacecraft, which was launched on November 19, 2010 on the DoD STP-S26 mission, carries three instruments developed in joint collaboration by NASA GSFC and the US Naval Academy: PISA, TTI, and MINI-ME.1,2 As part of a rapid-development, low-cost instrument design and fabrication program, these instruments were a perfect match for FASTSAT, which was designed and built in less than one year. These instruments, while independently developed, provide a collaborative view of important processes in the upper atmosphere relating to solar and energetic particle input, atmospheric response, and ion outflow. PISA measures in-situ irregularities in electron number density, TTI provides limb measurements of the atomic oxygen temperature profile with altitude, and MINI-ME provides a unique look at ion populations by a remote sensing technique involving neutral atom imaging. Together with other instruments and payloads on STP-S26 such as the NSF RAX mission, FalconSat-5, and NanoSail-D (launched as a tertiary payload from FASTSAT), these instruments provide a valuable “constellation of opportunity” for following the flow of energy and charged and neutral particles through the upper atmosphere. Together, and for a small fraction of the price of a major mission, these spacecraft will measure the energetic electrons impacting the upper atmosphere, the ions leaving it, and the large-scale plasma and neutral response to these energy inputs. The result will be a new model for maximizing scientific return from multiple small, distributed payloads as secondary payloads on a larger launch vehicle.

Fractal Loop Heat Pipe Heat Flux and Operational Performance Testing

This study investigates heat flux performance for a LHP that includes a fractal based evaporator design. The prototype Fractal Loop Heat Pipe (FLHP) was designed and manufactured by Mikros Manufacturing Inc. and validation tested at NASA Goddard Space Flight Center’s Thermal Engineering Branch laboratory. Heat input to the FLHP was supplied via cartridge heaters mounted in a copper block. The copper heater block was placed in intimate contact with the evaporator. The evaporator had a circular cross-sectional area of 0.877 cm2 . Twice distilled, deionized water was used as the working fluid. Thermal performance data was obtained for three different Condenser/Subcooler temperature combinations under degassed conditions (Psat = 25.3 kPa at 22°C). The FLHP demonstrated successful start-ups in each of the test cases performed. Test results show that the highest heat flux demonstrated was 75 W/cm2 .

Investigation of Gravitational Effects On Fractal Loop Heat Pipe Performance

This study investigates inclined operational performance for a LHP that includes a fractal based evaporator design. The prototype Fractal Loop Heat Pipe was designed and manufactured by Mikros Manufacturing Inc. and validation tested at NASA Goddard Space Flight Center’s Thermal Engineering Branch laboratory. Heat input to the FLHP was supplied via cartridge heaters mounted in a copper block. The copper heater block was placed in intimate contact with the evaporator. The evaporator had a circular cross-sectional area of 0.88 cm. Twice distilled, deionized water was used as the working fluid. Thermal performance data was obtained for five different inclinations (evaporator elevation ranging 30° above and below the condenser/subcooler bottom) using a single condenser/subcooler temperature combination (i.e., 10°C/5°C) under degassed conditions (Psat of 25.3 kPa at 22°C). The FLHP demonstrated successful start-ups in each of the test cases performed and had a maximum heat flux of 89 W/cm with the evaporator elevated 15° above the condenser bottom.

Fractal Loop Heat Pipe Performance Comparisons of a Soda Lime Glass and Compressed Carbon Foam Wick

This study compares heat flux performance of a Loop Heat Pipe (LHP) wick structure fabricated from compressed carbon foam with that of a wick structure fabricated from sintered soda lime glass. Each wick was used in an LHP containing a fractal based evaporator. The Fractal Loop Heat Pipe (FLHP) was designed and manufactured by Mikros Manufacturing Inc. The compressed carbon foam wick structure was manufactured by ERG Aerospace Inc., and machined to specifications comparable to that of the initial soda lime glass wick structure. Machining of the compressed foam as well as performance testing was conducted at the United States Naval Academy. Performance testing with the sintered soda lime glass wick structures was conducted at NASA Goddard Space Flight Center. Heat input for both wick structures was supplied via cartridge heaters mounted in a copper block. The copper heater block was placed in contact with the FLHP evaporator which had a circular cross-sectional area of 0.88 cm(sup 2). Twice distilled, deionized water was used as the working fluid in both sets of experiments. Thermal performance data was obtained for three different Condenser/Subcooler temperatures under degassed conditions. Both wicks demonstrated comparable heat flux performance with a maximum of 75 W/cm observed for the soda lime glass wick and 70 W /cm(sup 2) for the compressed carbon foam wick.

Project Phoenix: Design Considerations for a Student Liquid Bi-Propellant Rocket

To reestablish an enduring rocketry program at the United States Naval Academy, a group of midshipmen of the aerospace engineering department began designing a sounding rocket to reach an apogee of just over 6 miles (~10km). The student team from 2010 chose a liquid bipropellant motor for the propulsion system, a semimonocoque structure, and a parachute recovery system. The following year, subsequent students completed designs for the rocket subsystems. Seven midshipmen from the class of 2011 modified the designs and began construction of the various rocket subsystems as their senior design project. Based upon a starting rocket mass of 475 lbs (215 kg), a requirement was derived for 2200 lbf (9790 N) of thrust at a burn time of 17 seconds to reach the altitude objective. When designing components for the rocket, the team also had to take into consideration cost effective material and component selection as well as the fabrication capabilities of the machine shop at the Naval Academy. This paper serves as an overview of the rocket design.