ALBERT

Description: No abstract available

Description: An unsteady Reynolds averaged Navier-Stokes analysis loosely coupled with a comprehensive rotorcraft code is presented for a second-generation active-twist rotor. High fidelity Navier-Stokes results for three configurations: an isolated rotor, a rotor with fuselage, and a rotor with fuselage mounted in a wind tunnel, are compared to lifting-line theory based comprehensive rotorcraft code calculations and wind tunnel data. Results indicate that CFD/CSD predictions of flapwise bending moments are in good agreement with wind tunnel measurements for configurations with a fuselage, and that modeling the wind tunnel environment does not significantly enhance computed results. Actuated rotor results for the rotor with fuselage configuration are also validated for predictions of vibratory blade loads and fixed-system vibratory loads. Varying levels of agreement with wind tunnel measurements are observed for blade vibratory loads, depending on the load component (flap, lag, or torsion) and the harmonic being examined. Predicted trends in fixed-system vibratory loads are in good agreement with wind tunnel measurements.

Description: An unsteady Reynolds averaged Navier-Stokes analysis loosely coupled with a comprehensive rotorcraft code for blade trim and aeroelastic effects is presented for a second-generation Active-Twist Rotor. Mesh and temporal sensitives of computed airloads are evaluated. In the final paper, computed airloads will be compared with wind tunnel data for the Active-Twist Rotor test that is currently underway.

Description: Numerical predictions of the acoustic characteristics of an Active Twist Rotor (ATR), using two methods to compute the rotor blade aerodynamics and elastic blade motion are compared to experimental data from a wind tunnel test in the NASA Langley Transonic Dynamics Tunnel (TDT) in 2000. The first method, a loosely coupled iterative method, utilizes the Computational Fluid Dynamics (CFD) code OVERFLOW 2 and the Computational Structural Dynamics (CSD) code CAMRAD II. The second method utilizes the CAMRAD II free-wake model only. The harmonic active-twist control to the main rotor blade system is identified with three parameters - harmonic actuation frequency, actuation amplitude, and control phase angle. The resulting aerodynamics and blade motion data from the two methods are then used in the acoustics code PSU-WOPWOP to predict acoustic pressure on a spherical array of equally spaced observers surrounding the rotor. This spherical distribution of pressure is used to compute the sound power level representing baseline and actuated conditions. Sound power levels for three categories of noise are defined as - blade-vortex interaction sound power level (BVIPWL), low frequency sound power level (LFPWL), and overall sound power level, OAPWL. Comparisons with measured data indicate the CFD/CSD analysis successfully captures the trends in sound power levels and the effects of active-twist control at advance ratios of 0.14 and 0.17. The free-wake model predictions show inconsistent sound power levels relative to the trends in the experimental and CFD data. This paper presents the first ever comparison between CFD/CSD acoustic predictions for an active-twist rotor and experimental measurements.

Description: Projection moir interferometry (PMI) was employed to measure blade deflections during a hover test of a generic model-scale rotor in the NASA Langley 14x22 subsonic wind tunnel s hover facility. PMI was one of several optical measurement techniques tasked to acquire deflection and flow visualization data for a rotor at several distinct heights above a ground plane. Two of the main objectives of this test were to demonstrate that multiple optical measurement techniques can be used simultaneously to acquire data and to identify and address deficiencies in the techniques. Several PMI-specific technical challenges needed to be addressed during the test and in post-processing of the data. These challenges included developing an efficient and accurate calibration method for an extremely large (65 inch) height range; automating the analysis of the large amount of data acquired during the test; and developing a method to determinate the absolute displacement of rotor blades without a required anchor point measurement. The results indicate that the use of a single-camera/single-projector approach for the large height range reduced the accuracy of the PMI system compared to PMI systems designed for smaller height ranges. The lack of the anchor point measurement (due to a technical issue with one of the other measurement techniques) limited the ability of the PMI system to correctly measure blade displacements to only one of the three rotor heights tested. The new calibration technique reduced the data required by 80 percent while new post-processing algorithms successfully automated the process of locating rotor blades in images, determining the blade quarter chord location, and calculating the blade root and blade tip heights above the ground plane.

Description: An analytical study was conducted examining the feasibility of a swashplateless rotor controlled through two trailing edge flaps (TEF), where the cyclic and collective controls were provided by separate TEFs. This analysis included a parametric study examining the impact of various design parameters on TEF deflections. Blade pitch bearing stiffness; blade pitch index; and flap chord, span, location, and control function of the inboard and outboard flaps were systematically varied on a utility-class rotorcraft trimmed in steady level flight. Gradient-based optimizations minimizing flap deflections were performed to identify single- and two-TEF swashplateless rotor designs. Steady, forward and turning flight analyses suggest that a two-TEF swashplateless rotor where the outboard flap provides cyclic control and inboard flap provides collective control can reduce TEF deflection requirements without a significant impact on power, compared to a single-TEF swashplateless rotor design.

Description: The results of a computational study examining the effects of non-harmonic active-twist control on blade-vortex interaction (BVI) noise for the Apache Active Twist Rotor are presented. Rotor aeroelastic behavior was modeled using the Comprehensive Analytical Model of Rotorcraft Aerodynamics and Dynamics code and the rotor noise was predicted using the acoustics code PSU-WOPWOP. The application of non-harmonic active-twist inputs to the main rotor blade system comprised three parameters: azimuthal location to start actuation, azimuthal duration of actuation, and magnitude of actuation. The acoustic analysis was conducted for a single low-speed flight condition of advance ratio mu=0.14 and shaft angle-of-attack, a(sub s)=+6deg. BVI noise levels were predicted on a flat plane of observers located 1.1 rotor diameters beneath the rotor. The results indicate significant reductions of up to 10dB in BVI noise using a starting azimuthal location for actuation of 90?, an azimuthal duration of actuation of 90deg, and an actuation magnitude of +1.5 ft-lb.

Description: The Ares I-X Flight Test Vehicle (FTV), launched in October 2009, carried with it over 243 buffet verification pressure sensors and was one of the most heavily instrumented launch vehicle flight tests. This flight test represented a unique opportunity for NASA and its partners to compare the wind-tunnel derived buffet environment with that measured during the flight of Ares I-X. It is necessary to define the launch vehicle buffet loads to ensure that structural components and vehicle subsystems possess adequate strength, stress, and fatigue margins when the vehicle structural dynamic response to buffet forcing functions are considered. Ares I-X buffet forcing functions were obtained via wind-tunnel testing of a rigid buffet model (RBM) instrumented with hundreds of unsteady pressure transducers designed to measure the buffet environment across the desired frequency range. This paper discusses the comparison of RBM and FTV buffet environments, including fluctuating pressure coefficient and normalized sectional buffet forcing function root-mean-square magnitudes, frequency content of power-spectral density functions, and force magnitudes of an alternating flow phenomena. Comparison of wind-tunnel model and flight test vehicle buffet environments show very good agreement with root-mean-square magnitudes of buffet forcing functions at the majority of vehicle stations. Spectra proved a challenge to compare because of different wind-tunnel and flight test conditions and data acquisition rates. However, meaningful and promising comparisons of buffet spectra are presented. Lastly, the buffet loads resulting from the transition of subsonic separated flow to supersonic attached flow were significantly over-predicted by wind-tunnel results.

Description: This NESC assessment examined the accuracy of estimating buffet loads on in-line launch vehicles without booster attachments using sparse unsteady pressure measurements. The buffet loads computed using sparse sensor data were compared with estimates derived using measurements with much higher spatial resolution. The current method for estimating launch vehicle buffet loads is through wind tunnel testing of models with approximately 400 unsteady pressure transducers. Even with this relatively large number of sensors, the coverage can be insufficient to provide reliable integrated unsteady loads on vehicles. In general, sparse sensor spacing requires the use of coherence-length-based corrections in the azimuthal and axial directions to integrate the unsteady pressures and obtain reasonable estimates of the buffet loads. Coherence corrections have been used to estimate buffet loads for a variety of launch vehicles with the assumption methodology results in reasonably conservative loads. For the Space Launch System (SLS), the first estimates of buffet loads exceeded the limits of the vehicle structure, so additional tests with higher sensor density were conducted to better define the buffet loads and possibly avoid expensive modifications to the vehicle design. Without the additional tests and improvements to the coherence-length analysis methods, there would have been significant impacts to the vehicle weight, cost, and schedule. If the load estimates turn out to be too low, there is significant risk of structural failure of the vehicle. This assessment used a combination of unsteady pressure-sensitive paint (uPSP), unsteady pressure transducers, and a dynamic force and moment balance to investigate the integration schemes used with limited unsteady pressure data by comparing them with direct integration of extremely dense fluctuating pressure measurements. An outfall of the assessment was to evaluate the potential of using the emerging uPSP technique in a production test environment for future launch vehicles. The results show that modifications to the current technique can improve the accuracy of buffet estimates. More importantly, the uPSP worked remarkably well and, with improvements to the frequency response, sensitivity, and productivity, will provide an enhanced method for measuring wind tunnel buffet forcing functions (BFFs).

Description: Increasing demand for improved capabilities of small unmanned aircraft systems (sUAS) has generated interest in improving the design techniques for these vehicles. sUAS have typically been designed using iterative methods with multiple prototypes, but advancements in aircraft design software will make it possible to generate conceptual designs of very small VTOL aircraft with reduced hardware prototyping. This paper describes a research effort to generate a conceptual design of an approximately 6-lb quadcopter using the NASA rotorcraft design software NDARC. Wind tunnel and hover test data are used to validate and refine the conceptual design results. The effects of parametric design variations on vehicle scale are shown. The design study described herein shows that the NDARC software, which was designed for full-scale rotorcraft, can be used to design and evaluate sUAS vehicles.