Browse Topic: Drag
RPM-controlled hexacopters offer mechanical simplicity and inherent redundancy, but are unable to re-trim under all failure cases in forward flight. This paper investigates the use of reverse-enabled rotors as a means of expanding the attainable trim envelope and improving fault tolerance in RPM-controlled hexacopters. Isolated rotor experiments are conducted to characterize thrust and torque behavior under forward and reverse rotation, providing validation data for aerodynamic modeling. A blade-element-based model implemented in the Rensselaer Multicopter Analysis Code (RMAC) is then used to perform comprehensive trim analyses for a 1200-lb-class hexacopter in hover and in cruise at the best-range speed of 65 kts. Post-failure trim solutions are evaluated for four configurations, including edge-first and vertex-first orientations with different rotor spin directions. Results show that enabling reverse rotation allows trim recovery for all single-rotor failure cases in cruise, including aft-rotor failures that are not trimmable with conventional RPM-controlled rotors. A systematic comparison of peak rotor torque, peak rotor power, and total aircraft power reveals that failure severity is governed primarily by yaw moment deficits arising from the combined loss of hub torque and aerodynamic drag. Among the configurations examined, the edge-first configuration with a counter-clockwise spinning rotor 1 exhibits the lowest rotor torque and rotor power requirements, post-failure.
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).
This study examines the ability of a large (1200 lb gross weight) hexacopter with collective pitch controlled rotors to tolerate single motor failure. The hexacopter is considered in various orientations, and the vehicle is trimmed with one motor inoperative (OMI). Unlike RPM-controlled hexacopters, which were trimmable but uncontrollable in hover, and were untrimmable in cruise with an aft-rotor failure; with pitch-control the hexacopter is controllable in hover as well as trimmable for failure of any rotor in cruise (including an aft rotor failure). The study examines how pitch controls, and thrust are redistributed amongst the operational rotors, post-failure, for the different hexacopter orientations. For each case, the maximum thrust and torque increases on any individual rotor, and the total power increase, post-failure is examined. It is found that the hardest to trim cases are those where the hub torque and the hub drag induced yaw moment of the failed rotor add, and fault compensation for these cases usually comes at a high cost in terms of torque penalty and power requirement (necessitating use of more powerful motors and batteries). The results from the study indicated that operating the pitch-controlled hexacopter in an edge-first configuration with a clockwise spinning rotor 1 is the best choice from a fault tolerance perspective.
The paper presents a general framework for building an aeromechanic model in FLIGHTLAB, suitable for high fidelity, pilot-in-the-loop simulator. The focus is on aerodynamic modeling of AW609 tiltrotor in Airplane Mode flight regime. The framework can be extended to helicopter and conversion modes with additional considerations for rotors-airframe aerodynamic interference. It can also be adapted to different tiltrotor geometries, with some adjustments depending on their peculiarities. The model uses Blade Element Theory loads evaluation of lifting surfaces, corrected with tabulated distributed loads to tune FLIGHTLAB predictions against high-fidelity aerodynamic references. Bluff bodies are modeled using force and moment tabulated data. Verification was conducted against reference data in wind tunnel mode and against flight data in trim analysis. The proposed method allowed to match lift distribution on slender bodies, as well as lift and drag integral loads, with aerodynamic references. Trim analysis has shown satisfying accordance with experimental data for all the comparison parameters. However, rotor aerodynamic modeling is still being investigated to improve correlations in torque prediction. Additionally, installation effects, such as those related to the interaction between the wing and fuselage with the rotors, are under examination to address biases in flap motion evaluation.
This paper carries out experimental investigation of propeller and wing interactions under various geometric variations such as the horizontal and vertical distance between the propeller axis and the leading edge of the wing under different angle of attack conditions for a half wing setup for a wing made of symmetric airfoil. Rotor and wing performance is measured using independent six-component load cells. Through this study it is identified that for a wing made of symmetric airfoil optimal aerodynamic performance is significantly influenced by the position of the propeller. Positioning the propeller near the leading edge (x/c = 0.25) and on the negative side of the y-axis (y/c = −0.75) yields the best lift-to-drag ratios and enhanced lift, particularly in the moderate α range (4°–6°). Forward movement of the propeller along the x-axis (towards x/c = 0.75 or 1.00) increases drag and adversely affects performance.
Current paper summarizes a correlation study of two flow solvers (CREATETE-AV Helios and Simcenter STAR-CCM+), routinely used at Sikorsky, with multiple model-scale wind-tunnel tests. The Helios modeling approach was aiming for a high-fidelity accurate simulation, whereas the STAR-CCM+ modeling approach was aiming for a fast turn-around time with reasonable solution accuracy with a relatively coarse mesh and simplifications. The two solvers generally agreed well with the test data within reasonable accuracy and captured the airloads and flowfield trends. The calculations presented herein show the impact of the turbulence model on component loads, the aerodynamic interactions among components, and the effect of transition modeling on rotor performance. The Reynolds-Averaged Navier-Stokes CFD model generally delayed separation and resulted in lower drag. By modeling the airframe supporting structure in CFD simulations, an improvement on correlation for inflow on the propeller plane was shown. Additionally, improvements in the rotor system L/De correlation were realized by including a turbulence-transition model, which reduces the rotor drag.
The UH-60A slowed rotor test campaign carried out at the 40- by 80-Foot Wind Tunnel at the U.S. Air Force's National Full-Scale Aerodynamics Complex (NFAC) provided valuable information of a classical helicopter rotor blades operating at very high advance ratios. This paper aims to show the correlation of the RCAS and HOST comprehensive analysis (CA) tools with respect to several experimental campaign cases. Particularly the influence of the rotor aerodynamic performance as a function of the advance ratio and the collective angle is studied. The influence of the shank drag modeling is observed and its importance to obtain accurate results is highlighted. The RCAS and HOST simulations are capable of reproducing the rotor performance trends observed in the test campaign. Furthermore, the correlation of RCAS and HOST with respect to the measured rotor loads data is studied for the advance rations of 0.4, 0.5 and 0.7 at iso-thrust coefficient conditions. The aerodynamic loads and the structural loads are analyzed by means of the sectional loads analysis and polar maps. Finally some conclusions are provided highlighting the correlation of the RCAS and HOST results as well as the impact of the advance ratio on the behavior of the blades.
This study presents computational analyses of coaxial rotor hub flows and validation against experimental data obtained from the fifth Rotor Hub Flow Prediction Workshop. Experiments were conducted in a 12-inch diameter water tunnel at Pennsylvania State Applied Research Laboratory, employing tomographic particle-image velocimetry (Tomo-PIV) and precise hub drag measurements. Three CFD codes (UMD Mercury, CREATETM-AV Helios, and OVERFLOW) utilizing hybrid Reynolds-Averaged Navier-Stokes (RANS) / Large Eddy Simulation (LES) modeling based on Spalart–Allmaras turbulence model, were applied to replicate and analyze hub flows. Counter-rotating coaxial rotor hubs under free-air condition was simulated as the simplest case and the hub drags are compared between the three CFD codes. The full water tunnel configuration, consisting of two hubs, a fairing, and shafts, was also simulated and compared to experimental results, with a focus on hub drag, wake velocity fields, and turbulence quantities. Results demonstrated that the computational frameworks effectively captured key flow physics, although some discrepancies in drag harmonics, wake velocity and turbulence intensity magnitudes were observed. Additionally, the study highlighted the impact of rotor hub geometry and installation of sail-fairing on drag and wake structures. These findings contribute to improve computational predictions, essential for designing high-speed rotor hub configurations.
A computational study is conducted on a coaxial rotor hub and sail fairing configuration to analyze hub surface forces and the characteristics of its downstream wake. The flow conditions and grids are based on experimental tests performed at the Penn State Applied Research Lab (ARL) Water Tunnel at a baseline Reynolds number. Grid development for the rotor hubs and sail fairing is done using Pointwise v18.04R1 and Chimera Grid Tools (version 2.2). Simulations are performed using NASA's OVERFLOW2.4b Reynolds Averaged Navier-Stokes solver. The drag forces on the rotor hubs are computed and compared to standalone drag data to analyze the effects of interactional aerodynamics. Flow features, frequency content and Reynolds stresses of the wake are analyzed. Frequency content and Reynolds stresses show clear spatial bias. The anisotropy of the Reynolds stresses is computed and used to determine the character of the wake turbulence.
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.
In the realm of transitioning eVTOL aircraft, hindrance may be placed on performance in each of the two flight modes due to the existence of apparatuses or devices intended wholly for the other mode. For example, the presence of wings will normally reduce hover endurance due to their weight, and the use of a plurality of exposed lift-propellers - for hover stability and control - can lower flight speed and range in airplane mode because of the excess drag. It would seem, then, that transitioning eVTOL aircraft are generally poor performers in any mode when compared to their dedicated, single-mode cousins. This paper explores another possibility, of substantial performance improvement when the devices or their use become elements augmenting performance in the other mode - or cross-modally. Through an example dual-propeller aircraft, several cross-modal elements - including phenomena like the fan-in-wing effect and the inverse of Custer's channel-wing effect - are identified and their merits expounded.
Within this paper redundancy concepts on electric propulsion systems - consisting of electrical sources, inverters, electrical machines, gearboxes and drag generation units - are discussed. In a first steps different possible concepts are explained. In a general section considerations on the possible concepts are made, with a special focus on the design of the inverters, electrical machines and gearboxes. Advantages and disadvantages are shown and therefore some general assumptions on possible applications discussed. Later, two engineering examples for the concepts of shared drag generation unit and shared electrical machines with inverters are shown. The functionality is shown on measurement examples and experiences made during the design and testing phases are given. Finally, a new concept to reduce the risk of failure propagation in multi-wound motors is shown and discussed.
ABSTRACT
ABSTRACT
There are a large number of curves and slopes in the mountainous areas. Unreasonable acceleration and deceleration in these areas will increase the burden of the brake system and the fuel consumption of the vehicle. The main purpose of this paper is to introduce a speed planning and promotion system for commercial vehicles in mountainous areas. The wind, slope, curve, engine brake, and rolling resistances are analyzed to establish the thermal model of the brake system. Based on the thermal model, the safe speed of the brake system is acquired. The maximum safe speed on the turning section is generated by the vehicle dynamic model. And the economic speed is calculated according to the fuel consumption model. The planning speed is provided based on these models. This system can guide the driver to handle the vehicle speed more reasonably. According to the simulation, compared to cruise control, speed planning can save fuel consumption at a mean value of 9.13% in typical mountainous areas. The field test of a typical commercial vehicle shows that this system can increase fuel efficiency by 4.26% compared to an experienced driver during a journey in a mountainous area.
The force, torque, and energy methods of measurement are all in common use and should yield the same test results. Effects of steering, traction, and non steady-state tire operations are excluded from the recommended practice because they are still in the research stage. Methods of correcting laboratory data to road conditions are being developed.
A computational investigation of aerodynamic drag of coaxial rotor hubs is performed and compared with test data from a prior experiment. The counter-rotating coaxial hub model considered is based on a rotor design developed by AVX Aircraft Company. Component-level contributions to overall hub drag are quantified by building up the rotor from a bare shaft to complete hubs and control systems and measuring the drag as components are added. Fifteen total configurations are considered with CFD and compared to the experimental drag measurements of ten configurations from prior wind tunnel test data. The drag of each hub configuration is presented relative to a baseline configuration featuring the complete coaxial rotor hub and control system without the blade fairings. The CFD results and experimental data verify that the total hub drag is reduced by 25%-29% by incorporating the blade fairings. Nearly half of the remaining drag is from the mast and pitch links, while the blade grip contributes another 25% of the overall drag. The CFD results also show high level of interference drag (9.95%) which includes aerodynamic interference between mast/split plate, upper/lower swash plates, and swash plate/split plate.
Harold F. Pitcairn, American aviation and Autogiro pioneer, died from a single gunshot wound to the head in the late evening hours of April 23, 1960 at the age of 62 after a gala evening at which he presided over a celebration attended by more than 450 guests for his brother's Raymond's 75th birthday. Initially labelled a suicide by the press, Pitcairn's widow Clara declared that "she never wanted to hear another word about the tragedy", while friends and friendly local authorities made the argument, duly reported by Frank Kingston Smith in Legacy of Wings, his devotional Pitcairn biography (subsidized by the Pitcairn family), that the death was accidental because "there was no note, no indication of depression or unhappiness" and "the police investigation disclosed that two shots had been fired; one had penetrated the ceiling directly over the desk in the first floor study, another had struck Pitcairn in the eye" and that "the next morning it was discovered the semi-automatic pistol was defective: when cocked, it had a supersensitive "hair trigger," and it had a faulty disconnector so that it would fire more than one shot at a time, a condition known as "doubling."" The Pitcairn families, prominent and powerful, prevailed upon the local authorities to declare the death accidental and Kingston Smith's 1981account became the de facto authoritative story of the death of Harold F. Pitcairn. With the perspective, however, of six decades, it appears far more likely that Pitcairn's death was a suicide for reasons that were not readily evident, minimized, unappreciated or deliberately ignored at the time to craft a result that met the needs of Clara Pitcairn and her surviving family. These included the fact that while the claim was made that Pitcairn was making his nightly rounds to check on the estate’s ground-level windows (and had been doing so since the Lindbergh kidnapping in 1932), he actually died at his desk; that those in the house only reported a single shot; the 1907 Savage pistol had no reputation for a hair-trigger, and had not evidenced such a flaw in almost three decades of Pitcairn's nightly ritual; that even though Pitcairn had been assured that his almost-decade-long lawsuit against the United States government for Patent infringement of his Autogiro patents was going well, he was concerned about the impact this lawsuit was having on his aged associates who had been called to give depositions and he had voiced the sentiment that "if he had known that he would have to sue the government, he would not have gone into the Autogiro business"; that the lawsuit, itself intended as a vindication of Pitcairn's contribution to aviation was dragging on and would reach its first legal conclusion in 1967, and not finally conclude upon appeal until 1977; and most importantly, those who deny suicide and point to Pitcairn’s state-of-mind, have failed to take into account when the death occurred or ready evidence of his 'state of mind' To fail to see the tragic end of Harold F. Pitcairn is to forget that 29 years and one day earlier, he had been recognized for "the greatest achievement in aeronautics or astronautics in America, with respect to improving the performance, efficiency, and safety of air or space vehicles, the value of which has been thoroughly demonstrated by actual use during the preceding year." The memory of that day on the White House back lawn with the President was the high point of his life even as Pitcairn prepared to celebrate his older brother's achievements. The evidence, when marshalled and documented, conclusively points to suicide - a death of an American aviation pioneer before his contributions were vindicated in the largest patent infringement judgement against the United States in history. To fail to see the tragic end of Harold F. Pitcairn is to forget that 29 years earlier, he had been recognized for "the greatest achievement in aeronautics or astronautics in America".
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.
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.
The present study focuses on morphing of the inboard section of a helicopter rotor blade to improve aerodynamic performance in high-speed flight with significant reverse flow. Starting with the SC325218 airfoil, CFD simulations show that morphing only its lower surface, aft of the spar, to a modified ellipse geometry results in a significant reduction of reverse flow drag on the retreating side while limiting aerodynamic penalties in normal flow conditions on the advancing side. The study develops a morphing structure concept that transitions between the SC325218 airfoil and the derivative geometry while being able to carry aerodynamic loads. Key to this concept is the design of a morphing cellular lattice in the mid-chord region (25-65% chord), operating in collaboration with a specialized twosegment lower surface skin, and actuation ribs connected by spanwise stringers which also support aerodynamic loads in the aft-chord section. The cellular lattice's geometry is determined through optimization and the lattice is successful in morphing to the target geometry for high-speed operation. Parametric variations in lattice modulus, lower surface compliant skin modulus, and actuation strain are conducted toward minimizing actuation load and peak lattice strains, and key insight into the operational aspects of this morphing system are developed.
The rotor hub asembly is a primary contributor to rotorcraft parasite drag. Reducing hub drag is one mandatory step to enabling future high - sped conventional and compound rotorcraft. The importance of high - Reynolds number testing of rotor hub flows is emphasized by realizing that high - Reynolds number turbulent coherent structures remain strong for long distances downstream up to the long - age wake where they interact with the empennage and tail. Basic research conducted through the Vertical Lift Research Center of Excellence (VLRCOE) at Pen State's water tunnel facilities has provided unique high Reynolds - scale data of rotor hub wakes, providing new data for physical understanding and validation of computa tional fluid dynamics (CFD) methods. A first rot or hub flow prediction workshop was held in June 2016; the present paper focuses on 'blind comparison results' between experimental data and CFD analyses that were part of the second rotor hub flow prediction workshop at Pen State VLRCOE in May 2018. The hub workshops are the result of a collaboration between Pen State and Georgia Tech VLRCOEs with support from the National Rotorcraft Technology Center (NRTC), including participation from academia, industry, and government. They are compelling examples of productive exchange between experimental and computational efforts that advance the community's knowledge about these complex flows that are relevant to efficient and safe vertical lift.
This paper describes the development of a suction and oscillatory blowing (SAOB) active flow control (AFC) system, that was aimed towards achieving drag-reduction in a full-scale rotorcraft flight-test environment. The experimental work presented here includes benchtop characterization of the AFC actuators, development and testing of the AFC system using a full-scale two-dimensional airfoil at Tel Aviv University, and full scale testing of a UH-60 External Stores Support System (ESSS) in the CCDC AvMC Aviation Development Directorate 7- by 10-ft wind tunnel up to 140 knots. The airfoil tests composed of two stages, steady suction through holes on the trailing edge and the addition of pulsed blowing using SAOB actuators. The steady suction testing focused on identifying the optimal suction locations, while the SAOB experiments validated the suitability of the chosen actuator arrays to obtain the desired drag reduction at reasonable energetic cost. Tests for both steady suction and SAOB actuators resulted in 15-30 percent drag reduction compared to the clean baseline on the two-dimensional ESSS airfoil. Results from testing the full-scale ESSS identified the complexity of this configuration and the difficulty of transitioning directly from 2D component testing to three-dimensional testing of actual flight hardware. It became evident that by placing SAOB actuators only on the ESSS wing section, only 2-3 percent drag reduction was achievable. Further study is on-going to better understand aerodynamic interactions and develop a path towards more robust drag reduction.
High speed rotorcraft transmissions are subject to load-independent power losses consisting of drag and pumping loss. Tightly conforming shrouds enclosing the transmission gears are often incorporated to reduce the drag component of the total load independent losses. However, tightly conforming axial shrouding can result in an increase in the pumping loss component. Quantifying the pumping loss of shrouded gear transmissions has been the subject of many studies. This study presents a new approach for estimating pumping loss based on the concept of swept volume borrowed from the positive displacement pump and compressor industry. In this study, pumping loss of shrouded gear transmissions is considered to be related to the swept volume of the gear sets and the downstream flow resistance created by the shroud clearances. The drag loss and pumping loss of a spur gear pair have been determined through testing using the NASA Glenn Research Center Gear Windage Test Facility. The results from this testing have been compared to theoretical results using the formulations presented in this study. Good correlation exist between the test pumping power loss and the predicted pumping power loss for tightly conforming axial shroud configurations.
No abstract. Part of Introduction: Helicopter rotor hubs are geometrically complex components that experience a wide rage of aerodynamic behaviors and flow physics. This includes strong unsteadiness, large amounts of separation, laminar-turbulent transition, and interactional aerodynamic behaviors (Ref. 1). At high forward flight speeds (high advance ratios), the parasitic drag of the hub accounts for O(30 percent) of the total power required to fly (Ref. 1). A common method for characterizing this contribution is the hub drag factor, Kf e, which correlates the flat-plate area of the hub with the helicopter gross weight and functions as a technology factor (Ref. 2). In a recent assessment of needs for future vertical lift systems, Ormiston suggested that the hub drag factor needs to be reduced from the current state of the art of Kf e = 0:5 down to a value 0.2 (Ref. 3).
In order to extend the boundaries of helicopter performance and increase forward-flight speed, it is necessary to reduce the drag on the rotor hub, which can account for as much as 30% of the total parasite drag on the helicopter. Currently, there is limited experimental data available to predict the drag force on new hub configurations. The purpose of this testing is to create a database of lift and drag at various angles of attack to aid in hub design and hub drag prediction. Testing was conducted in the 12 inch-diameter water tunnel at ARL Penn State on four shapes - DBLN 526, 4:1 Ellipse, 3.25:1 Rectangle, and a new Optimized Cambered Shape (OCS) designed at UT Knoxville. Load cell data for lift and drag were obtained for angles of attack from approximately -5 degrees to 5 degrees. Drag data were also calculated using PIV velocity fields. Results are plotted and tabulated for use in future hub drag prediction toolsets.
Items per page:
50
1 – 50 of 710