Wing F. Ng

Improving Control System Effectiveness for Ducted Fan VTOL UAVs Operating in Crosswinds

A valuable new resource being developed for today’s soldier are small ducted fan VTOL UAVs. Although beneficial in many ways, a major operational problem of ducted fan vehicles is precise control when flying in crosswinds and turbulent conditions in general. There are two significant, inherent issues associated with ducted fan control in crosswinds; 1) lateral momentum drag and 2) a duct stabilizing torque which resist tipping into the wind. For vehicles designed for flight over large ranges in angle of attack (i.e. “transition to high-speed cruise”), trim over the flight envelope is strongly affected by the duct moment characteristics. Poor CG placement can lead to control saturation at intermediate flight speeds. These are characteristics of all ducted fan vehicles designed to hover as well as transition to horizontal flight. Techsburg, in collaboration with AVID LLC. and Honeywell Labs, has recently completed a research program to investigate and test new aerodynamic control devices to improve the handling of small ducted fan UAVs in crosswinds. This project consisted of both computational analysis and modeling of the vehicle aerodynamics and control system, as well as wind tunnel experiments using a powered ducted fan VTOL UAV model. This work leveraged methodologies and testing resources developed as part of a larger DARPAfunded VTOL UAV program. The research focused on a better understanding of the baseline vehicle and control vane aerodynamics, and selection and aerodynamic modeling and testing of new VTOL UAV auxiliary control device concepts. The analytical results are then used as input into an existing control system modeling framework, which in turn can be used to predict the vehicle’s dynamic response to turbulence. In addition to an improved flight control system, insights into establishing disturbance rejection criteria for VTOL UAVs are sought for use in future vehicle development efforts. This paper focuses on describing the baseline vehicle aerodynamics in a crosswind, as well as characteristics and design considerations for traditional control vanes used for this type of vehicle. Some limited, initial results from analysis of auxiliary control devices are presented as well.

High‐frequency temperature and pressure probe for unsteady compressible flows

A 3‐mm‐diam, dual hot‐wire aspiring probe is described which can simultaneously measure total temperature and total pressure in an unsteady high‐speed gas flow. The probe consists of two coplanar constant temperature hot wires at different overheat ratios operated in a 1.5‐mm‐diam channel with a choked exit. Thus, the constant Mach number flow by the wires is influenced only by free‐stream total temperature and pressure. The probe design is a compromise between the conflicting requirements of spatial resolution, frequency response, and angular sensitivity. The dc temperature accuracy of the probe is about 1% while the resolution is 0.3%. Frequency response of the present design is dc to 20 kHz.

Experimental Demonstration of Active Flow Control to Reduce Unsteady Stator-Rotor Interaction

An experimental investigation is conducted to reduce the unsteady stator-rotor interaction in a turbofan simulator using active flow control. The fan rotor of a 1/14-scale turbofan propulsion simulator is subjected to circumferentially periodic inlet flow distortions, generated by four stators that support a centerbody in the inlet mounted onto the simulator. These wakes are reenergized by injecting air from the trailing edge of each stator through discrete blowing holes. The flow rate through each blowing hole is controlled by an individual microelectro-mechanical system based microvalve. The microvalve actuation signal voltage is generated by a proportional-integral-derivative controller and is a function of the wake velocity defect. To determine the successful reenergizing of the wakes, far-field sound pressure level at the blade passing frequency without and with blowing is measured in an anechoic chamber. The active control experiments are performed for two simulator speeds of 29,500 and 40,000 rpm. In addition, the feasibility and advantage of active control is demonstrated by the ability of the system to respond to changes in the inlet flow velocity.

Effectiveness of a Serpentine Inlet Duct Flow Control Technique at Design and Off-Design Simulated Flight Conditions

An experimental investigation was conducted in a static ground test facility to determine the effectiveness of a serpentine inlet duct active flow control technique for two simulated flight conditions. The experiments used a scaled model of a compact, diffusing, serpentine, engine inlet duct developed by Lockheed Martin with a flow control technique using air injection through microjets at 1% of the inlet mass flow rate. The experimental results, in the form of total pressure measurements at the exit of the inlet, were used to predict the stability of a compression system through a parallel compressor model. The inlet duct was tested at cruise condition and angle of attack flight cases to determine the change in inlet performance due to flow control at different flight conditions. The experiments were run at an inlet throat Mach number of 0.55 and a resulting Reynolds number, based on the hydraulic diameter at the inlet throat, of 1.76* 10 5 . For both of the flight conditions tested, the flow control technique was found to reduce inlet distortion at the exit of the inlet by as much as 70% while increasing total pressure recovery by as much as 2%. The inlet total pressure profile was input in a parallel compressor model to predict the changes in stability margin of a compression system due to flow control for design and off-design flight conditions. Without flow control, both cases show a reduction in stability margin of 70%. With the addition of flow control, each case was able to recover a significant portion (up to 55%) of the undistorted stability margin. This flow control technique has improved the operating range of a compression system as compared to the same inlet duct without flow control.

Experimental Investigation of a Supersonic Shear Layer with Slot Injection of Helium

We focuse on mixing in a shear layer developed from helium slot injection into a parallel supersonic airstream and compare the results to those of previous slot-injection tests. In addition to short-duration schlieren and shadowgraph photography, concentration, pitot, cone-static, and stagnation temperature measurements are presented to document the development of the mixing layer.

Experimental comparison of two hot-wire techniques in supersonic flow

The performance of two constant-temperature normal hot-wire techniques for resolution of turbulent mass flux and local stagnation temperature in a supersonic flow is examined. The first technique used a single wire and the rapid scanning of multiple overheat ratios. Time averages of the signals at all overheats were used to separate the mean and rms mass flux, stagnation temperature, and their cross-correlation. The second technique used a dual-wire probe with each wire operated at different overheat ratios, giving instantaneous mass flux and stagnation temperature. In spite of a small separation distance (0.18 mm) between the wires in the dual-wire probe and high correlation between their signals, the rms mass flux inferred from the dual-wire technique was a factor of two higher than that from the single-wire technique. A consistency check based on data from one of the wires indicated that the dual-wire method produced results that were too high.

Aerodynamic/Aeroacoustic testing in Anechoic Closed Test Sections of Low-speed Wind Tunnels

This paper describes new anechoic closed test sections in Japan Aerospace Exploration Agency (JAXA) and Virginia Tech (VT) not only for an aeroacoustic but also for an aerodynamic testing ability in Low-speed wind tunnels. The anechoic closed test section with Kevlar wall is an innovative concept and had been originally developed at VT. It was later applied to JAXA’s 2m x 2m wind tunnel. By using a high-lift device model, the JAXAs new test section was evaluated and validated acoustically by comparing to VT anechoic test results. Moreover, the aerodynamic characteristics in the new test section were also evaluated by comparing to results of the same model in JAXAs closed hard-wall test section. New wall interference correction procedure is proposed for the Kevlar wall test section, and it showed very good agreement with well-known corrected hard-wall results. This anechoic test section is useful and a promising tool for both aerodynamic and aeroacoustic testing.

Experimental Measurements and Modeling of the Effects of Large-Scale Freestream Turbulence on Heat Transfer

The influence of freestream turbulence representative of the flow downstream of a modern gas turbine combustor and first stage vane on turbine blade heat transfer has been measured and analytically modeled in a linear, transonic turbine cascade. High intensity, large length-scale freestream turbulence was generated using a passive turbulence-generating grid to simulate the turbulence generated in modern combustors after passing through the first stage vane row. The grid produced freestream turbulence with intensity of approximately 10–12% and an integral length scale of 2 cm (Λx /c = 0.15) near the entrance of the cascade passages. Mean heat transfer results with high turbulence showed an increase in heat transfer coefficient over the baseline low turbulence case of approximately 8% on the suction surface of the blade, with increases on the pressure surface of approximately 17%. Time-resolved surface heat transfer and passage velocity measurements demonstrate strong coherence in velocity and heat flux at a frequency correlating with the most energetic eddies in the turbulence flow field (the integral length-scale). An analytical model was developed to predict increases in surface heat transfer due to freestream turbulence based on local measurements of turbulent velocity fluctuations and length-scale. The model was shown to predict measured increases in heat flux on both blade surfaces in the current data. The model also successfully predicted the increases in heat transfer measured in other work in the literature, encompassing different geometries (flat plate, cylinder, turbine vane and turbine blade) and boundary layer conditions.Copyright

ACTIVE FLOW CONTROL TO REDUCE FAN BLADE VIBRATION AND NOISE

An experimental investigation is conducted to reduce the unsteady stator-rotor interaction in a turbofan simulator using active flow control. The fan rotor of a l/lCscale turbofan propulsion simulator is subjected to circumferentially periodic inlet flow distortions, generated by four stators that support a centerbody in the inlet mounted onto the simulator. These wakes are re-energized by injecting air from the trailing edge of each stator through discrete blowing holes. The flow rate through each blowing hole is controlled by individual MEMS (Micro-Electra-Mechanical-System) based microvalve. The microvalve actuation signal voltage is generated by a PID (Proportional Integral and Derivative) controller and is a function of the wake velocity defect. Far-field Sound Pressure Level (SPL) at the Blade Passing Frequency (BPF) without and with blowing is measured in an anechoic chamber. The experiments are performed for two simulator speeds of 29,500 ‘pm and 40,000 rpm. Wake reenergization produces significant reductions in the BPF tone at both speeds. The sound power level at the BPF calculated from measured far-field directivity shows that source power is reduced by at least half The feasibility and advantage of active control is demonstrated by the ability of the system to respond to changes in the inlet flow velocity.

Novel Kevlar-Walled Wind Tunnel for Aeroacoustic Testing of a Landing Gear

which allowed aeroacoustic measurements to be carried out in the far field and in an environment with significantly less reflections. The model was a very faithful replica of the full-scale landing gear, designed to address the issues associated with low-fidelity models. A 63-element microphone phased array was used to locate the noise-source components of the landing gear from different streamwise positions, both in the near and far fields. The same landing-gear model was previously tested in the original hard-walled configuration of the tunnel with the same phased array mounted on the wall of the test section (i.e., near-field position). The new anechoic configuration of the Virginia Polytechnic Institute and State University wind tunnel offered a unique opportunity to directly compare data collected in hard-walled and semi-anechoic test sections, using the same landing-gear model and phased-array instrumentation. Through these tests, some of the limitations associated with testing in hard-walled wind tunnels were addressed. Nomenclature Cj = components of the steering vector d = distance from the array to the model f = frequency ffull-scale = full-scale frequency fmeasured = measured frequency k = wave number M = flow Mach number rj = distance traveled by an acoustic ray from the grid point with coordinates xn to array microphone j Rj = distance between the grid point with coordinates xn and array microphone j xn = coordinates of the grid point to which the array is being steered � = wavelength