Browse Topic: Computer simulation
Helicopter blades are often modeled as one-dimensional (1D) beams and considered to undergo medium-to-large deformations. The degree of nonlinearity that a beam theory can handle greatly affects prediction accuracy. In this work, quantitative evaluations are made for static and dynamic behavior of beams and blades using the classic moderately-large deformation beam (MLB) model as reference to a geometrically exact beam (GEB) model. A rotorcraft aeromechanics analysis framework is constructed to incorporate both beam models. The framework contains various solution procedures such as trim, blade response, loads, and vibrations while allowing external interface to high-fidelity computational fluid dynamics (CFD) analysis. A validation study is performed to examine the extent of the accuracy in the large deformation behavior of benchmark beam problems in static and dynamic conditions. Next, the HART (Higher-harmonic Aeroacoustic Rotor Test) II rotor is applied to evaluate the relative prediction accuracy of the beam models. Numerical simulations indicate that the MLB model shows poor correlation with the test data or fails to reach stable solutions as the level of loads and geometric twist angles become increased substantially. In case of HART II rotor, however, the MLB-based predictions show reasonable correlations with the measured data while indicating good agreements with the GEB results. The centrifugal stiffening by the rotor rotation causes limitations in the blade deformation which essentially lead to comparable solutions between the two beam results.
With recent advancements in the field of Advanced Air Mobility (AAM), including Electric Vertical Takeoff and Landing (eVTOL), Remotely Piloted Aircraft Systems (RPAS), and Unmanned Aerial System (UAS), it is beneficial to understand the impact of complex flow features on operations in urban and shipboard environments. Testing methods for studying these impacts, including simulated environments such as wind-tunnel flows and engineered equivalence tests, will need to be adapted to prepare for when the vehicles of interest are too large for the available testing facilities, and to permit low-cost alternatives for industry and government. This work demonstrates a development process that can be used to ensure the complex-flow-environment phenomena can be studied. First, this work illustrates the development of downdraft and turbulence flow types in a wind tunnel setting, and assesses the response of an M600 RPAS to these flows. Then, the same parameters are compared for a Mission Task Element (MTE) that was designed to challenge the RPAS in the vertical direction. Comparison of the two data sets provides insight into both the behaviour of RPAS in complex flows and the challenges of developing engineered equivalence tests.
This study numerically investigates the relationship between airspeed, drop height, and ground water coverage during helicopter-based aerial firefighting. With the effect of global warming and human activities the threat of forest fires has increased and finding optimal water dumping strategies for effective suppression is a crucial part of the firefighting operations. How varying airspeed and water drop height influence water dispersion and ground coverage has been analyzed utilizing numerical simulations with the VOF model in STAR-CCM+. Findings show that to maximize firefighting efficiency, balancing two contradicting phenomena is essential. These are, minimizing ineffective mist formation due to high drop height/high airspeed and fueling of the fire from rotor downwash due to low height/low airspeed passing by over the fire zone.
This paper presents findings from a joint computational-experimental venture that seeks to advance the physical understanding and validation-quality database for a model-scale generic tractor proprotor–wing system during the tiltrotor conversion maneuver. This study evaluates the interactions in a quasi-static manner for various proprotor tilt angles (θ) across the tiltrotor conversion maneuver. Independent experimental measurements of the wing and proprotor loads accompany synchronous wing surface pressure measurements along with stereoscopic particle image velocimetry flow field measurements at discrete spanwise locations. High-fidelity computational fluid dynamics simulations leverage the multi-disciplinary rotorcraft simulation tool CREATE™-AV Helios to assess the interactional aerodynamics of the proprotor–wing configuration across the tiltrotor conversion maneuver. Computational simulations use a newly implemented Helios module to trim to the experimental proprotor thrust. Validation of the computational model, along with a mesh refinement study, is shown via comparison of the integrated proprotor and wing loads as well as wing surface pressure distributions. The validated computational model is used to examine the proprotor-to-wing aerodynamic interactions further. For low proprotor tilt angles (0° ≤ θ ≤ 30°), interactions between the wing boundary layer and proprotor wake results in the breakdown of the wake along the wing chord. For moderate proprotor tilt angles (45° ≤ θ ≤ 60°), the interactions are dominated by the proprotor tip vortices curling around the wing leading edge. Approaching edgewise flight (75° ≤ θ ≤ 90°), the predominant proprotor–wing interactions stem from standing vortices on the wing upper surface, producing strong low-frequency responses on the wing. Overall, this work introduces new insights into the complex proprotor–wing aerodynamic interactions across the tiltrotor conversion.
Blade Vortex Interaction (BVI) noise primarily occurs in rotorcraft when tip vortices generated by the blades interact with other blades. When BVI noise occurs, it dominates at mid and high blade passing frequency harmonics. To mitigate BVI noise, we employ leading-edge serrations on the OLS rotor between 75% blade span and the tip. High-fidelity computational fluid dynamics simulations, using delayed detached eddy simulation, combined with an acoustic analogy, are conducted to analyze various leading-edge serration geometries with different serration height and wavelength parameters. The results show that rotor BVI noise is reduced by up to 5 dB at the rear of the vehicle when serrations are applied, with higher serration height-to-wavelength ratios proving more effective. The findings demonstrate that when vortices directly impinge on the rotor blades, the serrations disrupt the vortices and generate a fluctuating pressure field on the blade surface, leading to destructive phase interference at the far-field microphone locations.
The Dragonfly relocatable lander was selected as NASA's New Frontiers mission in 2019 to explore the organic-rich surface of Titan, Saturn's largest moon. The coaxial quadrotor vehicle will fly to multiple geologic sites covering a distance of over 50 miles near the Titan equator. At each site, Dragonfly will sample materials, determine the surface composition, and investigate how far prebiotic chemistry has progressed on Titan. Upon arrival, the lander will enter the Titan atmosphere protected inside an aeroshell, which will descend and decelerate with parachutes. At an altitude of approximately 1 km above the ground, the lander will separate from the backshell and perform a controlled transition to powered flight. Prior to separation from the backshell and after the heatshield has been ejected, the Preparation for Powered Flight (PPF) sequence will be initiated, which ensures the lander is in a safe and stable state for autonomous descent. A critical element of PPF is the de-spin maneuver, where diagonally opposing rotors rotate at maximum speed to reduce any residual angular momentum by creating a yaw moment in lander body axes. This paper presents high-fidelity computational fluid dynamics simulations of the Dragonfly rotorcraft lander during the PPF sequence. Aerodynamic performance predictions are compared with test data from the National Full-Scale Aerodynamics Complex to validate the simulations and build confidence in the PPF simulation results. Blade-resolved simulations capture the unsteady and complex flow behavior in Titan's dense, low-temperature atmospheric conditions during PPF. The results are analyzed, providing insight into aerodynamic performance and the aerodynamic moments critical for mission success.
Leveraging lessons learned from NASA's Ingenuity Mars helicopter and concepts such as the Mars Sample Recovery Helicopter, and Mars Science Helicopter has enabled partners at NASA's Jet Propulsion Laboratory (JPL), NASA Ames, and AeroVironment, Inc. to mature a hexacopter vehicle concept (Chopper) with the ability to support a wide range of mission scenarios. This work focuses on the critical aeronautics-related challenges encountered transitioning from an Ingenuity-size vehicle to a much larger vehicle (˜15 times the mass) and discusses engineering efforts to address these challenges. Critical upgrades include optimized airfoils, higher solidity blades, and higher fidelity computational models. Because multiple rotors are required to lift the heavier vehicle, increased understanding of the impact of rotor-to-rotor interactions is also necessary. Rotors have been designed that are tailored to more demanding missions and will be validated in a joint test campaign between the partners. While the Chopper concept will be utilized to illustrate these maturation efforts, the lessons learned are applicable to other heavier next generation Mars rotorcraft platforms also.
ABSTRACT Northrop Grumman has developed a software and hardware solution to provide enhanced 360 degree local situational awareness (LSA) to enable the warfighter with an overmatch capability on today’s modern battlefield. The architecture exploits technological gains in cameras, video processing, and video compression. The approach allows rapid comprehension of local and remote situational views presented with operational relevance for a ground combat platform or tactical wheeled platform crew. The 360 Degree LSA approach provides direct visualization of relative positioning of targets, threats, and lines of fire; and additionally offers common situational understanding / operational picture from the dismounted soldier to higher echelon commands. The approach provides prioritized information through LSA software to provide an enhanced view to the warfighter whereas the squad leader becomes an integral part of the crew with a view of the common operating picture (mounted) and additional sensors on tablet or handheld device (dismounted via wireless). The approach uses a platform agnostic form factor with components that can be selected and applied to legacy or new platforms based on their size, weight, power, and mission constraints.
ABSTRACT Rubber tracks are now extremely competitive for vehicles up to 50 tons and fully fielded on 39 ton vehicles. They represent the best of what technology can offer for tracked vehicles, in terms of high durability, performance and low life cycle cost. This is mainly attributed to the optimization through the five (5) technological tools described in this paper. Better from its numerous distinctive advantages, rubber tracks can be adapted to suit virtually any specific need. This ductile rubber track technology can be shaped to match today’s requirements, with the help of advanced rubber compounding and computer simulations.
WHY DO WE NEED SIMULATIONS? This paper is intended to provide a broad presentation of the simulation techniques focusing on transmission testing touching a bit on power train testing. Often, we do not have the engine or vehicle to run live proving ground tests on the transmission. By simulating the vehicle and engine, we reduce the overall development time of a new transmission design. For HEV transmissions, the battery may not be available. However, the customer may want to run durability tests on the HEV motor and/or the electronic control module for the HEV motor. What-if scenarios that were created using software simulators can be verified on the test stand using the real transmission. NVH applications may prefer to use an electric motor for engine simulation to reduce the engine noise level in the test cell so transmission noise is more easily discernable.
ABSTRACT The authors studied the effects of different types of armor on the performance of spin-torque microwave detectors (STMD). Working prototypes of novel nano-sized spintronic sensors of microwave radiation for battlefield anti-radar and wireless communications applications are being integrated into Sensor Enhanced Armor (SEA) and Multifunctional Armor (MFA) and tested in SEA-NDE Lab at TARDEC. The preliminary theoretical estimations have shown that STMD based on the spin-torque effect in magnetic tunnel junctions (MTJ), when placed in the external electromagnetic field of a microwave frequency, can work as diode detectors with the maximum theoretical sensitivity of 1000 V/W. These STNO detectors could be scaled to sub-micron size, are frequency-selective and tunable, and are tolerant to ionizing radiation. We studied the performance of a STMD in two different dynamical regimes of detector operation: in well-known traditional in-plane regime of STMD operation and in recently discovered novel out-of-plane regime.
Army rotorcraft operations demand precision and adaptability to navigate challenging terrain, respond to real-time mission requirements, and ensure time-to-target arrival. Navigating complex terrain, making real-time trajectory adjustments, and ensuring timely arrival at designated objectives while considering other problems are challenging. This paper focuses on the concept of 4D conformal pilot cueing that can facilitate a significant reduction in pilot workload. To enhance the rotorcraft operations with Army scenarios, a visual cueing method based on Tau Theory for obstacle avoidance is proposed so that the pilot can make a coordinated turn away from the obstacle and safely change the helicopter's trajectory to avoid the collision. To demonstrate the visual cueing method, desktop simulations are performed in Matlab/Simulink environment using simulated pilots.
Rotorcraft shipboard operations are risky and demand high piloting skills. Computational simulations are invaluable for pilot training and understanding of the complicated aerodynamic environment. Yet, simulations may not capture a realistic response of the rotorcraft due to simplified modeling of the interactional aerodynamics. Thus, improvements to modeling and simulation are required, and experimental data are needed that unveil the interactional aerodynamics and dynamics between the rotor and ship for computational validation efforts. In this study, an extensive experimental investigation of the ship-rotor dynamic interface problem was conducted to gain a general understanding of the interactional aerodynamics between a 1:100 wind-tunnel-scale NATO Generic Destroyer and a representative single main rotor, operating both stationary and dynamically in the vicinity of the landing deck. Data obtained through simultaneous measurements of rotor hub loads, ship deck surface pressures, and stereoscopic particle image velocimetry flow fields gave valuable insight into the highly coupled aerodynamic phenomena. Results showed that the rotor hub loads exhibited a high dependency on both the wind direction and also the position of the rotor relative to the landing deck. A bifurcation in the regions of high unsteady thrust was observed under certain wind conditions due to the ship airwake impacting different portions of the rotor disk. The inflow angles across the rotor disk were estimated and assessed through flow field measurement, revealing how the ship airwake altered the rotor inflow and wake structure that contributed to the unique change in rotor loads under various hovering conditions.
ABSTRACT Higher-order methods and adaptive mesh refinement (AMR) have demonstrated their added value to numerical simulation of a variety of aerodynamic problems, including rotorcraft flow. Within this paper, the implementation of a higher-order spatial WENO (Weighted Essentially Non-Oscillatory) method for flux reconstruction and an AMR method for grid refinement into a comprehensive and approved rotorcraft simulation framework is discussed. A significant increase in computational efficiency by factor 15 can be achieved using higher-order methods. AMR allows a considerable reduction of required overall grid cells, and thus computational effort, to represent the occurring flow phenomena with equivalent accuracy. The rotorcraft simulation framework is applied to simulate a highly complex compound helicopter geometry with a variety of challenging phenomena. Flight mechanic trim and elastic blades are taken into account for a physical representation of the flight state. The benefit of the AMR technique in combination with the higher-order methods for an isolated rotor in hover and a complete helicopter configuration in cruise flight is investigated. A robust application of the AMR implementation is demonstrated, which shows a reduction of computational cost by 50% at high advance ratio cases without a noticeable loss of accuracy.
This paper experimentally investigates direct effects of lightning strikes on flax fiber-reinforced polymers. Highcurrent artificial lightning strikes are conducted on coupon level to evaluate thermo-mechanical damage and to quantify the sufficiency of copper wire mesh as lightning strike protection (LSP). The dataset shall also serve for verification of prospected numerical simulation. The natural fiber flax, as a sustainable source of composite reinforcement, has been demonstrated to be suitable for semi-structural parts of rotorcraft. However, its low electrical and thermal conductivity requires a functional LSP layer for aviation applications. The test panels are investigated regarding their material combination, stacking sequence and level of LSP. Results show that two as well as three layers of 72 g/m2 copper mesh are not sufficient to withstand the standardized lightning current component A waveform of 200 kA. The high induced currents and low capability of energy dissipation leads to electro-explosion of metal and transient mechanical forces from shock waves causing mechanical damage on the test panels. Back surface-velocities increase with higher peak currents and higher level of protection results in lower damage. It is shown that a stacking of copper wire mesh results in less arc root dispersion.
This study addresses safety concerns within the rapidly evolving Electric Vertical Takeoff and Landing (eVTOL) aircraft domain, focusing on efficient tools to quantify uncertainties in lithium-ion battery behavior - a critical aspect of eVTOL. One major issue with quantifying uncertainty is the prohibitive computational cost associated with many queries of an expensive-to-evaluate computational model. This work employs three physics-based battery models models of varying fidelity and cost to estimate the mean and the variance of the selected quantities of interest through a multifidelity method to reduce the computation cost. By combining information from multiple cheaper, lower-fidelity models through the Multifidelity Monte Carlo method, we significantly reduce the number of high-fidelity samples required for a prescribed mean-squared error, consequently reducing computational costs down to a tractable level. The proposed methodology is applied to estimate the mean and the variance of the battery temperature and voltage, accounting for uncertainties in flight conditions and materials. The first example focuses on a 580-second flight and is benchmarked against a standard Monte Carlo sampling technique. Results indicate a notable fourfold speed-up using the Multifidelity Monte Carlo method compared to the standard Monte Carlo method for the same mean-squared error for the voltage estimate. To showcase the method's generality, the multifidelity method is then applied to a longer flight of 3580 seconds for estimating the mean and the variance and utilizing these statistics to approximately estimate the probability of the flight completion. This demonstrates the adaptability of the methodology to various power profiles and considered uncertainties, with potential extensions to any battery chemistry. In conclusion, the presented multifidelity method offers a robust approach to enhance eVTOL safety by efficiently estimating uncertainties in battery behavior.
This paper presents the results of a research and development (R&D) effort focused on fluid structure interactions between airframe structures and bladder type fuel tanks during a crash environment. During this R&D effort, fuel tank and surrounding structure crash impact tests were conducted using an innovative test configuration that allowed low-cost fabrication of test articles which represented several different design architectures. LS-DYNA models of the crash test article configurations were also developed and correlated with the tests data. Good correlation between the test data and LS-DYNA analysis results was achieved. The paper also includes recommendations for design of the airframe structures around the fuel tanks based on the fluid structure interaction insights gained from the crash tests and analyses.
Small multirotor vehicles, for example, designed for package delivery, are expected to operate in close proximity to populated areas, raising concerns about noise pollution. This study utilizes acoustic flight tests and computational modeling of an instrumented research hexacopter developed at Penn State to investigate noise generation during takeoff and landing maneuvers, considering varying flight path angles and vehicle speeds. Flight tests were conducted at Mid-State Regional Airport and corresponding predictions were made using the Penn State Noise Prediction System. The predicted vehicle states and noise levels are first validated against the flight test data. The validated model and flight test data are then utilized to study the noise emissions of the aircraft. Measurements and predictions of the acoustic characteristics of the vehicle are analyzed using conventional noise metrics, frequency content, and directivity features. Descent maneuvers are found to be noisier than climb maneuvers. Noise generation decreases with an increase in forward speed during both climb and descent and also with an increase in vertical speed during climb. However, during descent, noise generation increases with an increase in sink rate.
A joint experimental-computational research campaign is underway to develop physical understanding and a validation-quality database for a model-scale tractor propeller-wing system. Separate load measurements on the wing and propeller accompany wing surface pressure distributions and flow field measurements via stereoscopic particle image velocimetry (SPIV) at discrete wing spanwise locations for a range of static propeller tilt angles. The physical wind tunnel test is modeled using a high-fidelity computational approach (Helios). Computational simulations aid in assessing the influence of the wind tunnel facility effects and test support structure wake interference, as well as in reducing uncertainties in the physical experiments for use in computational validation. The behavior of the induced thrust and lift at a zero-degree wing angle of attack in the axial flow regime (cruise configuration) is correlated with flow field measurements, showing distinct differences between upwash and downwash fields produced by the propeller. Recirculation regions near the wing upper surface are identified for high propeller tilt angles as the propeller wake interacts with the wing boundary layer. Increased propeller tilt angles result in a decreased wing lift and increased propeller thrust in forward flight, resembling a rotor in ground effect. Results indicate ways to improve computational modeling and refine the physical wind tunnel testing. These will be applied to refine the database to further elucidate physical understanding of the propeller-wing interactional aerodynamics.
In the field of aerodynamics, there is a growing need for rapid load prediction in engineering applications. Surrogate modeling offers a promising solution, providing faster results compared to high-fidelity computational models. This study focuses on a Machine Learning (ML) framework tailored for surrogate modeling, specifically for integrated aerodynamic load predictions in aircraft design. Central to this framework is a Deep Neural Network (DNN) component capable of handling both steady-state and fluctuating aerodynamics. A key challenge for surrogate models lies in maintaining prediction accuracy, especially in scenarios involving nonlinear flow phenomena like flow separation and transonic shifts. To address these challenges, we introduce a two-step physics-state predictor that integrates an intermediate Convolutional Neural Network (CNN) component. This approach enhances the surrogate model's capability to accurately represent dynamic separated flows and other nonlinear patterns without relying on unrealistic user inputs. Results are presented for NACA0015 dynamic stall predictions for two different physics-state inputs.
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In this work, we present a lightweight pipeline for robust behavioral cloning of a human driver using end-to-end imitation learning. The proposed pipeline was employed to train and deploy three distinct driving behavior models onto a simulated vehicle. The training phase comprised of data collection, balancing, augmentation, preprocessing, and training a neural network, following which the trained model was deployed onto the ego vehicle to predict steering commands based on the feed from an onboard camera. A novel coupled control law was formulated to generate longitudinal control commands on the go based on the predicted steering angle and other parameters such as the actual speed of the ego vehicle and the prescribed constraints for speed and steering. We analyzed the computational efficiency of the pipeline and evaluated the robustness of the trained models through exhaustive experimentation during the deployment phase. We also compared our approach against state-of-the-art implementation in order to comment on its validity.
In November 2019, NASA completed the first wind tunnel test entry of the Multirotor Test Bed (MTB), a new test capability for advanced VTOL rotorcraft configurations. The MTB had been under development since 2017 when the need arose for an easily reconfigurable test stand for multirotor aircraft configurations. With the wide-ranging assortment of aircraft currently targeted at Urban Air Mobility and Unmanned Aircraft System applications, there is a need for validation data that will increase confidence in the computational modeling tools being used to develop these platforms. The MTB fills this need. This paper describes the key features of the MTB as well as its first wind tunnel test entry. A selection of results from the test is presented here, demonstrating the flexible configuration of the MTB and the types of data researchers can generate using this new test capability.
The present paper designs and validates a finite element bird model in order to develop a useful tool for the numerical simulation of an aeronautical bird strike event. The bird is simulated in Abaqus/Explicit environment using the Smoothed-Particle Hydrodynamics (SPH) technique. This formulation is a common approach to the problem, as noticed in many published works that investigate the bird strike problem using other solvers such as LS-Dyna. They are taken as a starting point of the present work in the definition of the bird model initial geometry and equation of state. A comparison and a correlation between some experimental tests and their simulations was conduct, in order to develop and validate the bird numerical model. In particular, impacts against rigid targets, such as plates and wedges, and against deformable ones are evaluated. The result is a validated SPH FE bird model, which can be involved in the next bird strike analysis being a reliable numerical tool.
In this paper, an experimental and numerical study of a rotor interacting with wing of three different aspect ratios at an advance ratio of 0.5 is described. Those three wings have the equal area and fixed at an angle of attack of 8 degrees. One wing is installed at 3 different vertical positions and three different horizontal locations to investigate the influence of the wing position on the interaction. The calibration and correction process of the measurement is described, and the results are compared to pre-test CFD simulations. Numerical simulations based on simplified rotor and wing-body models have been carried out ahead of the wind-tunnel testing. Due to the existence of rotor-test-stand and the rotorhub which are not included in the CFD simulations, the measured aerodynamic performance deviated from the CFD results remarkably. By applying angle-of-attack and drag offset corrections which remove the influence of the rotortest-stand, the wing performance is found in good agreement with the CFD results. Also, applying hub-correction which is based on measured hub-only aerodynamic force components, and the angle-of-attack correction, the corrected isolated rotor performance in high advance ratio agrees satisfactorily with the CFD data. It is found that due to the rotor/wing interaction, the overall lift-to-effective drag ratios decrease about 12% at the advance ratio of 0.5.
The use of formation flight to achieve aerodynamic benefit as applied to rotorcraft is, unlike its fixed-wing counterpart, an unproven principle. This document presents a proof-of-concept of rotorcraft formation flight through a numerical research study, supported by results from an independent wind-tunnel experiment. In both cases, two helicopters are placed in an echelon formation aligned on the advancing side of the main rotor, though they do not simulate directly comparable flight conditions. The vertical and lateral alignment is varied in order to observe the achievable reductions in main rotor power required during cruise flight. The wind-tunnel experiment data yields an estimated maximum total power reduction for the secondary aircraft of 24%, while the numerical models yield reductions between 20% and 34% dependent on flight velocity. Both experiments predict a higher potential for aerodynamic benefit than observed for fixed-wing formations, which is contributed to the asymmetric upwash profile in the rotor wake. Optimal lateral alignment of both experimental and numerical results is found to feature overlap of the rotor disk areas due to circular area effects. Experimental data shows an optimal vertical alignment of the secondary rotorcraft below the primary, due to wake displacement. This is not present in the numerical simulations as a result of the applied leader wake modeling.
Military rotorcraft engines operating in harsh environments routinely ingest large quantities of mineral dust, which can degrade components and ultimately reduce operability. Time off-wing for unscheduled maintenance is a costly burden, both financially and operationally. Rapidly predicting engine deterioration rates as a function of the mission presents an opportunity to optimise flow of supplies, better manage fleets, and perform safety risk assessments when dust loading is expected to be particularly high. In the current contribution, we present our ongoing efforts in this field with a new methodology for assessing the effectiveness of inertial particle separators and quantifying the changes they impart to the inbound dust. We demonstrate that both the concentration reduction and the modification to the particle size distribution can be made on the basis of a single independent variable- a generalised Stokes number for inertial particle separators- and a single performance parameter- the corrected separation efficiency. To develop these parameters we conduct numerical simulations of the flow through a generic axi-symmetric inertial particle separator, over a range of five mass flow rates, three scavenge mass flow rates, and 16 particle diameters. In addition to this, a framework is presented to enable an estimation of the dust concentration at the engine intake. This is achieved by correlating the total wake strength to an existing dust landing trial dataset. A coupled rotorcraft-engine model is then used to combined the two methodologies to investigate the influence of engine mass flow rate on dust ingestion rate. A weak non-linear relationship is observed, which arises due to the simultaneous increase in wake strength with engine mass flow rate as rotor power requirements increase. The additional dust stirred up by the stronger wake leads causes this non-linearity. Finally, we show that an improvement in separation efficiency caused by higher engine mass flow rate is far outweighed by the associated increase in dust loading in this condition.
A pilot-in-the-loop simulation environment aimed at increasing pilot visual cues without the need of expensive visualization hardware is presented. The proposed solution relies on Virtual Reality (VR) to enhance the pilot immersion in the simulated environment. The project is integrated in the development of the complete simulation framework FRAME-Sim, focused on simulating rotorcraft in early conceptual design stages, and therefore relying on physics-based multibody simulation of the rotorcraft flight dynamics and free/open source software. FRAME-Sim visual environments that are being used include products available to the market as well as homemade solutions developed to obtain the highest level of versatility during the simulation.
A method for developing and performing a rotorcraft flight simulation in a bluff body airwake is investigated and compared to flight test data. This method is broadly composed of two parts: an atmospheric boundary layer computational solution that is unsteady and time-accurate, and a UH-60 flight dynamics program that can be coupled with an airwake developed from the atmospheric boundary layer solution. The flight dynamics program is a MATLAB/SIMULINK implementation of the GENHEL model. The computational fluid dynamics airwake solution is developed to recreate the airwake present for the validation case flight tests that were used to collect the flight test data. The validation case f light test data was collected from a variety of station-keeping tasks performed in the wake of a hangar on land. Two kinds of flight simulations were performed: piloted simulations in a flight simulator facility and flight simulations performed using a pilot model. Comparisons of the power spectral density analyses of pilot control stick activity reveal that pilots in the simulated flight tests use a generally higher level of control activity energy than the pilots of the f light test validation case. Limited simulator flight test data showed sizeable differences in pilot control activity for the same simulated flight task flown by different pilots. Significant unknowns remain in comparing a simulated flight to a physical flight. The pilot model was not used to compare simulated flights to validation case flights, but was useful for comparing different flight simulation cases to each other as well as for analyzing sensitivity of flight simulations to positioning error within a simulated airwake. The pilot model flight simulations show that flight tasks performed in an airwake produced using a high-fidelity atmospheric boundary layer simulation require more pilot control input power than airwakes made using a comparable mean wind speed uniform inflow.
Low-order rotor broadband noise models require the boundary layer, displacement, and momentum thicknesses around the rotor's component airfoils to be provided. This data is often unavailable from prior literature, and may not be accurately computed (if at all) by simplified aerodynamics codes. To obtain the boundary layer thicknesses, this study builds on an integrated and robust meshing and solution approach using HAM2D, a Navier-Stokes computational code with advanced flow transition computations. This approach using HAM2D is already used to compute airfoil performance tables, and the computation of the boundary layer thicknesses imposes negligible additional overhead. The boundary layers computed by HAM2D are validated against prior data and wind tunnel PIV measurements conducted specifically for this study. In general, HAM2D shows good agreement with both experimental data sets, particularly with the PIV measurements conducted for this study, and where HAM2D diverges from historical data, so too do the PIV measurements. A few specific points of disagreement are noted, and future work will seek to reduce these inconsistencies.
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