Compact tunable microfluidic interferometer

We demonstrate a compact tunable filter based on a novel microfluidic single beam Mach-Zehnder interferometer. The optical path difference occurs during propagation across a fluid-air interface ( meniscus), the inherent mobility of which provides tunability. Optical losses are minimized by optimizing the meniscus shape through surface treatment. Optical spectra are compared to a 3D beam propagation method simulations and good agreement is found. Tunability, low insertion loss and strength of the resonance are well reproduced. The device performance displays a resonance depth of - 28 dB and insertion loss maintained at - 4 dB. (C) 2004 Optical Society of America.

Thermal and electrical currents in nanoscale electronic interferometers

We theoretically study thermal transport in an electronic interferometer comprising a parallel circuit of two quantum dots, each of which has a tunable single electronic state which are connected to two leads at different temperature. As a result of quantum interference, the heat current through one of the dots is in the opposite direction to the temperature gradient. An excess heat current flows through the other dot. Although locally, heat flows from cold to hot, globally the second law of thermodynamics is not violated because the entropy current associated with heat transfer through the whole device is still positive. The temperature gradient also induces a circulating electrical current, which makes the interferometer magnetically polarized.

Effects of hydrogen-air non-equilibrium chemistry on the performance of a model scramjet thrust nozzle

Two aspects of hydrogen-air non-equilibrium chemistry related to scramjets are nozzle freezing and a process called 'kinetic afterburning' which involves continuation of combustion after expansion in the nozzle. These effects were investigated numerically and experimentally with a model scramjet combustion chamber and thrust nozzle combination. The overall model length was 0.5m, while precombustion Mach numbers of 3.1 +/- 0.3 and precombustion temperatures ranging from 740K to 1,400K were involved. Nozzle freezing was investigated at precombustion pressures of 190kPa and higher, and it was found that the nozzle thrusts were within 6% of values obtained from finite rate numerical calculations, which were within 7% of equilibrium calculations. When precombustion pressures of 70kPa or less were used, kinetic afterburning was found to be partly responsible for thrust production, in both the numerical calculations and the experiments. Kinetic afterburning offers a means of extending the operating Mach number range of a fixed geometry scramjet.

Dynamic force balances for short-duration hypersonic testing facilities

Two force balance techniques for use in hypersonic impulse facilities are compared by measuring the drag force on a 30° semi-apex-angle blunt cone model in a hypersonic shock tunnel at a free stream Mach number of 5.75. An accelerometer-based balance and a stress-wave force balance were tested simultaneously on the same model to measure the drag force. It was found that drag force measurements could be made using both techniques in a flow with a 450-μ s test period. The measured drag forces compared well with the theoretical values estimated using Newtonian theory.

Microfluidic interferometer

We present a class of compact, fluid-based, interferometric, tunable optical components: the single-beam microfluidic Mach-Zender interferometer. Phase delay is achieved through light propagation across a fluid-air interface (meniscus). The effect of meniscus curvature on the device transmission is considered using the three-dimensional beam propagation method and shown to be an important device parameter. We engineer the meniscus curvature using monomer surface chemistry, rendering it flat, and find that the experimental response corresponds well with simulation. The device has a resonance at 1.3 mum with a 25 dB extinction ratio; the latter can be adjusted by shifting the meniscus position.

Coupled thermal, structural and vibrational analysis of a hypersonic engine for flight test

This paper describes a relatively simple and quick method for implementing aerodynamic heating models into a finite element code for non-linear transient thermal-structural and thermal-structural-vibrational analyses of a Mach 10 generic HyShot scramjet engine. The thermal-structural-vibrational response of the engine was studied for the descent trajectory from 60 to 26 km. Aerodynamic heating fluxes, as a function of spatial position and time for varying trajectory points, were implemented in the transient heat analysis. Additionally, the combined effect of varying dynamic pressure and thermal loads with altitude was considered. This aero-thermal-structural analysis capability was used to assess the temperature distribution, engine geometry distortion and yielding of the structural material due to aerodynamic heating during the descent trajectory, and for optimising the wall thickness, nose radius of leading edge, etc. of the engine intake. A structural vibration analysis was also performed following the aero-thermal-structural analysis to determine the changes in natural frequencies of the structural vibration modes that occur at the various temperatures associated with the descent trajectory. This analysis provides a unique and relatively simple design strategy for predicting and mitigating the thermal-structural-vibrational response of hypersonic engines. (C) 2006 Elsevier SAS. All rights reserved.

Flight data analysis of the HyShot 2 scramjet flight experiment

The development of scramjet propulsion for alternative launch and payload delivery capabilities has been composed largely of ground experiments for the last 40 years. With the goal of validating the use of short duration ground test facilities, a ballistic reentry vehicle experiment called HyShot was devised to achieve supersonic combustion in flight above Mach 7.5. It consisted of a double wedge intake and two back-to-back constant area combustors; one supplied with hydrogen fuel at an equivalence ratio of 0.34 and the other unfueled. Of the two flights conducted, HyShot 1 failed to reach the desired altitude due to booster failure, whereas HyShot 2 successfully accomplished both the desired trajectory and satisfactory scramjet operation. Postflight data analysis of HyShot 2 confirmed the presence of supersonic combustion during the approximately 3 s test window at altitudes between 35 and 29 km. Reasonable correlation between flight and some preflight shock tunnel tests was observed.

Scramjet lift, thrust and pitching-moment characteristics measured in a shock tunnel

Lift, pitching moment, and thrust/drag on a supersonic combustion ramjet were measured in the T4 free-piston shock tunnel using a three-component stress-wave force balance. The scramjet model was 0.567 m long and weighed approximately 6 kg. Combustion occurred at a nozzle-supply enthalpy of 3.3 MJ/kg and nozzle-supply pressure of 32 MPa at Mach 6.6 for equivalence ratios up to 1.4. The force coefficients varied approximately linearly with equivalence ratio. The location of the center of pressure changed by 10% of the chord of the model over the range of equivalence ratios tested. Lift and pitching-moment coefficients remained constant when the nozzle-supply enthalpy was increased to 4.9 MJ/kg at an equivalence ratio of 0.8, but the thrust coefficient decreased rapidly. When the nozzle-supply pressure was reduced at a nozzle-supply enthalpy of 3.3 MJ/kg and an equivalence ratio of 0.8, the combustion-generated increment of lift and thrust was maintained at 26 MPa, but disappeared at 16 MPa. Measured lift and thrust forces agreed well with calculations made using a simplified force prediction model, but the measured pitching moment substantially exceeded predictions. Choking occurred at nozzle-supply enthalpies of less than 3.0 MJ/kg with an equivalence ratio of 0.8. The tests failed to yield a positive thrust because of the skin-friction drag that accounted for up to 50% of the fuel-off drag.

Performance of a scramjet combustor with combined normal and tangential fuel injection

The performance of a scramjet combustor with combined normal and tangential injection was experimentally investigated. Experiments were performed on a 500-mm cylindrical scramjet combustor at a freestream Mach number of 4.5, a nozzle supply pressure of 35.8 MPa, and a nozzle supply enthalpy of 5.8 MJ/kg. Hydrogen fuel was injected normally through portholes to promote combustion and tangentially through a slot to reduce viscous drag. A series of fuel injectors were used to vary the proportion of tangential to normal fuel between 45 and 100%. Reductions in the viscous drag of up to 25% were observed with the greatest reductions occurring at the lowest total equivalence ratio tested for each injector. However, the average pressure produced by combustion with combined normal and tangential injection was approximately 50% less than that produced by normal injection alone. An analysis of the change in specific impulse of the scramjet combustor indicated that the best overall performance was produced by 100% normal injection.