Browse Topic: Pressure
An experimental investigation was conducted to explore the loads, acoustics, and tip vortex trajectories of coaxial counter-rotating (CCR) rotor with unequal upper and lower radii. The upper and lower rotor radii were tested both at the nominal radius of 1.108 m, and also with a lower rotor radius of 90% nominal radius, for a constant rotor speed of 1180 RPM and a constant inter-rotor spacing of z/R = 0.108. Rotors were torque balanced and tested for a range of upper rotor collective pitch from -2◦ to 10◦ . The power required for both CCR systems was within 0.9% for most trim conditions, and equal thrust was produced at upper rotor collectives of 6◦ and 8◦ (within 1.0%). At low loading conditions the unequal radii configuration produced more thrust for the same power due to a reduction in profile drag. The overall sound pressure level (OASPL) was lower for the CCR rotor with shortened lower rotor blades at all angles of elevation. Larger reductions in A-weighted OASPL(A) were observed, due to a larger contribution of broadband noise to the total OASPL(A).
The influence of ground, wall, and corner boundaries on multirotor vehicle performance was investigated through a series of controlled flight tests. Changes in rotor inflow profiles were represented by near-field rotor pressure measurements captured by a custom Kiel probe wake rake. Ground effect was characterized by reduced thrust and power requirements, primarily driven by the vehicle fuselage, which induced regions of reduced pressure and increased flow unsteadiness around the airframe. Operating near a wall boundary was found to restrict airflow into the portion of the rotor disk closest to the wall, leading to increased power requirements to maintain hover and a consequent reduction in performance. While vehicle orientation had minimal impact on overall rotor performance, it did influence local rotor inflow behavior near the wall, depending on the relative position of the interaction region formed with adjacent rotors. As the vehicle descends from the isolated wall effect into corner effect, created by the intersection of the wall and ground, an exchange between the dominating ground and wall effects is observed, with corresponding ground heights and wall distances identified.
An experimental investigation was conducted on a 1.108 m radius coaxial co-rotating (or stacked) rotor and a coaxial counter-rotating (CCR) rotor of identical geometry to compare the acoustics and loads of both rotor configurations in hover. The rotors were operated at a tip Mach number of 0.40, tip Reynolds number of 765,000 and an axial spacing of 1.55 chord lengths, and the index angle between the upper and lower blades of the stacked rotor was varied. The overall sound pressure level (OASPL) was significantly larger for the CCR rotor. For example, total rotor noise at -45◦ angle of elevation was 6 dB greater for the CCR rotor than the stacked rotor at 8◦ collective. These increases in OASPL were driven by large increases in tonal noise for the CCR rotor, of up to 10 dB higher than the stacked rotor at some angles of elevation. This was attributed to additional tonal noise occurring at harmonics of 2Nb/rev, due to vibratory loads from 2Nb/rev blade crossings. The results of the experimental study suggest that, in terms of reduced noise and increased hover efficiency, the stacked rotor is preferable to the CCR rotor for eVTOL vehicle configurations that implement compact or closely spaced designs.
ABSTRACT A full-scale Reynolds number water tunnel experiment was performed to generate a data set used to analyze the effects of helicopter rotor hub wake impingement on a canonical horizontal stabilizer. The experiment was designed and performed in the Pennsylvania State University Applied Research Laboratory Garfield Thomas Water Tunnel, where a 10.5 inch constant chord stabilizer was placed in the 48-inch diameter test section downstream of a 1/4 scale helicopter hub. Two rotor hubs were tested, a baseline configuration and a low-drag model. The stabilizer was mounted in the long-age wake. Lift, pitching moments, and unsteady pressures were measured on the horizontal stabilizer at a Reynolds number of 0:9x10⁶, 1:8x10⁶ and 2:7x10⁶, corresponding to hub diameter-based Reynolds numbers of 2:2x10⁶, 4:3x10⁶, 6:5x10⁶ and rotor advance ratios of 0.1, 0.2, and 0.3. The hub-wake interaction results were compared to a baseline airfoil test, which was performed without a hub upstream. Pressure sensors were used to evaluate wake unsteadiness impinging on the horizontal stabilizer. The horizontal stabilizer in clean flow exhibited lift and pitching moment in agreement with XFOIL predictions. With the low-drag hub upstream it measured lift fluctuations at a frequency of 2/rev, 4/rev, 8/rev and 12/rev. Downstream velocity and pressure fluctuations of 2/rev 4/rev and notably 6/rev were measured with the baseline hub upstream. Drag reduction on the low-drag hub was measured to be >25% compared to the baseline hub at full-scale Reynolds number. Both drag and wake harmonics measured at the hub and downstream on the stabilizer were found to be dependent on the upstream hub geometry. Pressure frequencies taken on the horizontal stabilizer yielded similar results and were consistent with those measured via the force balance.
Installation effects of the Volocopter 2-X beam structures are studied by performing high-fidelity CFD simulations of a single and three-rotor configurations in hover. The studied cases are compared with simulations without airframe to investigate the installation effects. In addition, the noise emission of the configurations is simulated by using a Ffowcs Williams-Hawkings based CAA code. Scattering effects are also included by using a BEM code. The rotors are simulated at an identical RPM and are placed in their mounting position. Furthermore, an additional setup with individual rotor RPMs is simulated for the three-rotor configuration. The installation mainly affects the rotor wake, thrust and pressure fluctuations on the rotor, while the integral aerodynamic quantities remain almost unchanged. This resulted in additional oscillations in the acoustic pressure signal. Overall, the installation increases the OSPL by about 1.5 dB, but has a greater effect on the 3-20 harmonics. The simulation data were compared with the Volocopter measurements and showed good agreement. For the three-rotor configurations, the rotor-rotor interactions were found to be dominant for the aerodynamic and acoustic performance, while the installation effects only locally influenced the noise emission by 3-4dB.
Airfoil optimization for rotor blades is a critical endeavor aimed at enhancing aerodynamic performance and reducing noise. This paper employs a Kriging surrogate model combined with a multi-objective genetic algorithm to optimize thrust, power, and broadband noise. Three airfoil parameterization methods including ParFoil, PARSEC, and CST are compared when used to generate various airfoil shapes for the surrogate model and optimization process. We utilize low-fidelity aerodynamic tools such as XFOIL and blade element momentum theory for aerodynamics. In addition, acoustic modeling is conducted using Lee's wall pressure spectrum model alongside Amiet's trailing-edge noise model. The paper focuses on small-scale rotor configurations, specifically an ideally twisted rotor using the NACA 0012 airfoil and a modified XV-15 blade. Both blades are used as baseline models for hover optimization. The optimization of the ideally twisted rotor across various parameterization methods demonstrates a significant reduction in A-weighted overall sound pressure level by approximately 4.0 dBA. The primary contributors to this noise reduction are identified as a decrease in the chordwise pressure gradient and wall shear stress, which are key factors that contribute to the wall pressure spectrum and resulting trailing-edge noise. Similarly, the optimization of the modified XV-15 blade using the ParFoil method achieves a notable decrease in noise of 3.5 dBA. Further analysis is extended to axial flight conditions, with axial velocities of 5.0, 10.0, and 20.0 m/s, to simulate vertical climb. The results show that these hover-optimized blades are capable of sustaining noise decreases of 3.0-4.0 dBA compared to the baseline, provided that the thrust coefficient is maintained through adjustments in collective pitch.
ABSTRACT The paper investigates the influence of a dynamic active twist control on BVI noise in descent flight for a Mach-scaled Bo 105 model rotor. Therefore, a numerical study has been carried out with DLR's aeromechanics rotor code S4. In addition, sound pressure levels have been computed with DLR's acoustic code APSIM. The intensity of BVI noise is determined by different vortex parameters as well as the blade-vortex miss-distance and the BVI location. In this paper, all parameters that influence the intensity of the radiated BVI noise are analyzed. The analysis is carried out for a rotor shaft angle of attack of αᴿₒ = 4° for which the most intensive BVI noise was identified for the passive rotor blade. The study lays emphasis on the effect of different control amplitudes and phases but concentrates on a control frequency of the third multiple of the rotor rotational frequency for which highest reductions in BVI noise were identified.
This study examines the acoustics of a wing operating in the wake of a propeller. The propeller wing system is simulated at 24 knots cruise and 8° wing angle of attack. The propeller is simulated using an actuator line model, while the wing is simulated using two different turbulence models: a DDES turbulence model and a higher fidelity LES model. Chordwise compact loads, on-wing pressure surfaces, and pressure surfaces at distances of 2.34% and 10% thickness around the wing surface are used as inputs to PSU-WOPWOP to predict noise at an observer below the wing. Using on-wing surface pressures, the LES broadband noise predictions are 13.5 dB higher than DDES. Chordwise compact loads result in lower noise predictions than on-wing surface pressures, by 11.3 dB for LES and 2.3 dB for DDES. Using off-body pressure surfaces, DDES results remain similar to noise predictions from on-wing pressure surfaces, but with LES the broadband noise predictions are about 2.5 dB lower.
Ground effect analytical models in the literature, such as the Pressure Potential Based Finite State Model (PPBFSM) or Velocity Potential Based Finite State Model (VPBFSM) have been developed to study an isolated rotor in full ground effect. These models use the mass source distributions to account for the ground plane. Also, these models consider that the exit pressure of those distributions is equal to the pressure exerted from the rotor at the ground plane. However, PPBFSM and VPBFSM do not satisfy the non-penetration of flow boundary condition at the ground plane. This paper develops a new ground effect model using the VPBFSM that considers the non-penetration of flow boundary condition at the ground plane by varying ground rotor size and using optimization to find the strength of the mass source distributions. Additionally, it captures the flow at any point below the disk which is missing in the previous study of ground effect using VPBFSM. The developed model is applied to the R-50 unmanned helicopter rotor using its geometric and aerodynamic data from the literature. The results show a good correlation compared to the Hayden model. Additionally, the present model exhibits a similar trend of rotor inflow in ground effect at different rotor heights above ground compared to the PPBFSM.
Tailsitter configurations that operate in both fixed and rotary wing flight modes are typically capable of generating large control forces and moments, making them inherently capable of rapid transitions and aggressive maneuvers. However, harnessing these capabilities requires feedback control strategies that can effectively estimate the non-linear aerodynamics loads involved to successfully exploit them. This paper describes initial steps in combining an onboard flow sensing strategy with a data-driven approach to estimating inflight air loads. A neural network is trained to use measurements from a multi-hole probe to predict the output from a set of pressure sensors embedded in a wing section undergoing a series of pitch motions in a wind tunnel. We hypothesize that this limited context of emulating a sensor network represents a focused and compartmentalized approach to applying emerging data-driven techniques to challenging aeronautical problems. We compare estimation results from a set of neural networks with varying input configurations to assess the feasibility of our approach and the significance of different sensing modalities on overall performance. Current results show that a gated recurrent network (GRU) trained with unsteady pressure measurements was able to predict the chordwise pressure distribution on a pitching NACA 2412 airfoil using probe measurements, reproducing the transient and non-linear effects observed in our dataset.
ABSTRACT
This slash document collects general reference material related to gaseous oxygen system flow requirements and sizing calculations. This document will assist oxygen system equipment designers and operators to establish systems and equipment requirements. The document consists of charts, tables, system schematics, system requirements, and sample calculations for system sizing.
Computations were performed to assess the effect of fluidically-oscillating jets on a ROBIN-mod7 helicopter fuselage. The simulations utilize previously experimentally validated methodologies that rely on a new boundary condition formulation at the actuator throats, based on phase-averaged flow variables, which obviates the need to resolve the internal cavities simultaneously with the outer flow. Predictions of the base flow past the helicopter fuselage were validated against experimental and computational data available in the literature. The fluidic oscillator characteristics were then evaluated at different scales and pressure ratios, and invariant quantities were identified. In the flow control evaluation, flow separation was significantly reduced and, in some cases, suppressed. However, drag reduction was not obtained, indicating the sensitivity of the actuation location and operating conditions to the vehicle design and flight orientation.
Acceptance of close-proximity eVTOL aircraft depends on public perception that they are quiet in relation to the environment. When they are audible, their influence on community soundscapes must not be intrusive. To solve for that, measurement of "quiet" is needed using a physiological model of hearing rather than a sound pressure measurement intended for entirely different applications. While we are starting with baseline noise metrics using earlier methods, we propose a cooperative effort to coordinate development, testing and education about techniques that correlate better with perception of quiet. Two major differences are proposed between current practice and future methods: measure vehicle loudness rather than sound pressure, and measure operations with respect to ambient rather than to meet an arbitrary level for a given land use. These are not new concepts, but require computational resources that until recently were too expensive to be practical.
Acoustic characteristics of two- and four-bladed rotors with a 1.108 m radius in hover were measured experimentally at rotor speeds up to 1200 RPM and tip Mach number up to 0.41. The Rotorcraft Comprehensive Analysis System (RCAS) using a viscous vortex particle method (VVPM) coupled with PSU-WOPWOP were used to simulate the acoustic characteristics of the same rotors. Simulations were also conducted for a stacked rotor over a range of azimuthal spacings with zero axial spacing. PSU-WOPWOP predictions included thickness, loading, and broadband noise. Both experiments and simulations showed that broadband noise was the dominant contributor to A-weighted sound pressure level at a distance of 8.25 radii from the rotor center. Simulations showed a sharp increase in the broadband and overall noise at higher thrust, but the experiments did not show such an increase. Higher harmonic components of the tonal noise were nearly as large as the blade passage frequency noise in the experiments, but were underpredicted in the simulations by up to 30 dB. Simulations incorporating wind showed closer agreement in the higher harmonic tonal components. The unweighted overall sound pressure level in the stacked rotor simulations showed 5 dB to 10 dB variations with azimuthal spacing, but the A-weighted tonal and overall SPL was insensitive to azimuthal spacing.
High fidelity code-to-code comparisons have been made between the University of Glasgow HMB3 code and the HPCMP CREATE™-AV Helios code under The Technical Cooperation Program collaboration project, Next Generation Rotor Blade Design. The comparisons are made for two model-scale rotors - Langley baseline (LBL) rotor and Pressure Sensitive Paint (PSP) rotor. Hover and forward flight performance results are compared against test data. For the LBL rotor, hover performance is in a good agreement between the test data and HMB3 results over a full range of CT. However, the comparison between the HMB3 and Helios results at a CT of 0.0084 shows the difference in Figure of Merit (FM) by approximately 2 counts (2.2-3.2%). In forward flight, the HMB3 and Helios performance results overpredict the test data at the low advance ratios but improve the predictions at the high advance ratios. At an advance ratio of 0.31, the code-to-code comparison indicated that the Helios torque was lower by 2.8-3.1% compared with the HMB3 torque. For the PSP rotor, the FM trend computed using the Helios-OVERFLOW code agrees well with the fully-turbulent test data. The Helios-OVERFLOW result shows that the fuselage interaction effect is favorable for rotor performance. For a code-to-code comparison, the HMB3 FM result shows about 3.0 counts (4.0%) lower than the HeliosOVERFLOW result at a CT/σ of 0.091. In forward flight, the HMB3 rotor result agrees well with the test data whereas the Helios rotor result underpredicts by 2-3 counts in CP/σ (3.5-5.7%) at the high advance ratios although it was significantly improved at the low advance ratio. Overall, the code-to-code comparisons are successfully conducted for both the LBL and PSP rotors.
The British Experimental Rotor Programme (BERP) tip design is well known for its superior performance for high speed flight. This paper revisits the BERP design by presenting a parameterized model of the planform design based on published data and investigating its performance on the Apache rotor blade. The underlying airfoil sections, HH02 and NACA64A006, are retained in a new Apache BERP-shape rotor blade. The performance of the Apache BERP-shape rotor blade is evaluated for hover and forward flight by using US Army CREATE™-AV Helios software and compared with the Apache baseline rotor blade. The results are presented in the form of rotor thrust, rotor torque, figure of merit, trim condition, sectional blade loading and pressure distribution. A visualization of complex vortex structures is provided to offer insight into the airflow characteristics at different flight conditions.
ABSTRACT Rotorcraft operations in arid environments can result in the ingestion of large quantities of dust particles into turboshaft engines, where they can melt and deposit on high pressure turbine nozzle guide vanes. This can result in reduced engine life-span and in worst case scenarios, in-flight engine failure. Predicting the extent and rate at which this damage occurs has proven difficult owing to the wide range of variables relating to the dust cloud, engine and most importantly, the properties of the particulate encountered. Whilst significant work has been carried out to model the particle deposition process for both volcanic ash and coal fly-ash, there is scarce similar work for the different types of mineral dusts rotorcraft encounter. In this contribution, we assess the suitability of two opposing numerical approaches for use in a generalised, reduced-order deposition model of individual mineral particles depositing on a vane. Both models are seen to be heavily reliant upon empirical inputs, be this the thermo-mechanical properties of the particles such as their yield strength, or currently unknown experimentally determined constants. An alternative approach is therefore proposed whereby the particle yield strength is correlated using existing relationships to the Vickers hardness of the grain, a property more amenable to empirical determination. The results obtained represent the current applicability limits of the two models based upon existing empirical data and thus highlight the need for further experimentation relating to both the thermo-mechanical properties and probabilities of adhesion for both individual mineral grains and mineral dust blends.
This paper presents an efficient prediction of coaxial rotor broadband noise, particularly trailing-edge noise. The method combines a newly developed iterative coaxial rotor BEMT, a viscous panel method, an empirical wall pressure spectrum, and Amiet's trailing-edge noise model. Aerodynamic data including the induced velocity and angle of attack on each rotor are calculated by the iterative BEMT. Then, turbulent boundary layer flows, such as the boundary layer thickness, skin friction coefficient, pressure gradient, etc., are computed by a viscous panel code, XFOIL. Based on these boundary layer parameters, the wall pressure spectrum near the trailing edge is computed by Lee's semi-empirical model. Finally, trailing-edge noise is predicted by Amiet’s model from the wall pressure spectrum. This method provides fast computations for aerodynamics and acoustics for coaxial rotors. Acoustic predictions can be performed for various design and operating conditions including the effect of rotor-to-rotor separation distance. In addition to the overall noise of the combined rotor system, each rotor's contributions to noise can be analyzed. A small-scaled untwisted rotor is selected to analyze aerodynamics and aeroacoustics. It is found that the noise contribution from each rotor is about the same at small separation distances. At large separation distances, the lower rotor generates higher noise levels than the upper rotor mainly due to the change in rotor distance with respect to the observer. The detailed boundary-layer flow properties are investigated on both rotors.
Auralization of unsteady helicopter flight operations is needed to better understand the impact of maneuvers on community noise. Previous source sound synthesis methods were based on interpolated data, which may lead to artifacts in generating sound for helicopter maneuvers where sound pressure directivity may change rapidly. In this paper, the source sound is synthesized at every time sample using a solution to the Ffowcs Williams-Hawkings Equation along the path of the emission angle between the source and a ground observer. The synthesized sound is then propagated to the ground observer for auralization. Since no interpolation is performed, maneuvers with rapidly changing sound pressure directivities may be more accurately synthesized and auralized. The framework for accomplishing this synthesis and auralization is described, which couples the Fundamental Rotorcraft Acoustic Modeling from Experiments, the second-generation Aircraft NOise Prediction Program, and the NASA Auralization Framework. Synthesis of a hovering rotor is presented to compare with previous synthesis methods. Two examples with aperiodic signals are then presented to demonstrate synthesis and auralization of unsteady rotor noise.
The purpose of this study is to determine if phase synchronization can be used to reduce the net radiated sound power from two rotors. Phase synchronization implies that the rotors have the same rotational speed with a fixed relative azimuthal blade position, or phase. The concept is evaluated both experimentally and numerically. Measurements of source directivity and thrust are initially compared with predictions to confirm that the model accurately captures the relevant trends. The model is then used to explore the design space and identify relevant parameters. Both experimental and numerical results show that the radiated sound power at the blade passage frequency can be reduced by appropriately controlling the relative azimuthal phase of the rotors. Vehicle level predictions are also provided for a notional octocopter, comparing two different modes of operation. Predictions show that phase synchronization can be used to achieve a 4-5 dB reduction of the sound pressure level at the blade passage frequency nearly everywhere on the ground plane beneath the vehicle.
This paper discusses an effort to determine useful scaling laws for predicting the sound pressure levels at various observer positions for small scale rotors in hover. Surveys of rotor thrust, rotor torque and near-field acoustics are acquired using a series of identically shaped two-bladed rotors ranging between 8 and 18 inches in diameter. Rotor speeds were varied between 50 and 90 revolutions per second which equate to blade tip Mach numbers as high as 0.56. Common sound metrics like overall sound pressure level, sound pressure level and blade-pass frequency harmonic noise were correlated with both dimensional variables (rotor thrust, torque, power and blade tip speed) and non-dimensional variables (thrust and power coefficients, Reynolds number, blade tip Mach number and figure of merit) pertaining to rotor performance. The findings revealed a new correlation parameter for predicting rotor noise based on the product of figure of merit with blade tip Mach number. It is postulated that an increase in rotor collective translates to an increase in loading noise. So long as the flow remains attached, the increased loading noise is expressed by an increase in figure of merit. This new correlation parameter is shown to provide a reasonable collapse of all rotor noise data. In order to account for the directivity of the sound field, a second order response surface model of this new correlation parameter was developed and validated.
Aerodynamic interactions between the rotor and the empennage can have a significant impact on steady and unsteady loads and often result in challenges in a rotorcraft design phase. In the present work, numerical analysis of rotor-empennage aerodynamic interactions were compared to full-scale flight test data with respect to steady and unsteady interactional aerodynamic effects. The flight tests provided loads for a low-empennage and a T-Tail configuration for various forward flight velocities. For the T-Tail configuration, additional pressure sensors provided validation data for steady and unsteady interaction effects. The numerical analysis was focused on an unsteady panel method, complemented by high-fidelity CFD/CSM-coupling results for a level flight state. Furthermore, a supplemental validation of the unsteady panel method was performed against an isolated wing-vortex interaction experiment. The flight test data revealed a strong asymmetry in mean empennage loads, which increases with forward flight velocity. The numerical analysis showed coherent results with a slight over-prediction in high-speed. The T-Tail configuration is furthermore subject to 3D effects between the vertical- and horizontal tail. These effects influence pressure and load-distributions on the T-Tail, which was captured by both numerical methods. The general characteristics of the unsteady pressure signatures were captured by both methods. The panel method showed slightly better representation of amplitude.
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