Search

Measurements in the wake of a free-flying full-scale helicopter in ground effect were performed for both quasi-steady and unsteady maneuvering flights using stereoscopic particle image velocimetry (PIV), a time-resolved background-oriented schlieren (BOS) setup, and an optical marker tracking technique. The systems were used in a complementary way to both visualize blade tip vortices in a large portion of the rotor wake and to capture spatially resolved wake velocity data close to the ground. The high sensitivity of the BOS system enabled the detection of vortices up to an age of ψ = 630°. Different instability mechanisms as long-wave, short-wave, and pairing instabilities were observed with varying intensity for different flight conditions. A quantitative analysis of vortex locations showed a periodic variation resulting from interactions between consecutive vortices that led to vortex pairing. Characteristics of the wake outwash close to the ground were investigated by means of averaged velocity fields. Different patterns such as wall jet, recirculation, and ground vortex flow were quantitatively analyzed and found to be in good agreement with previous model helicopter experiments. The instantaneous velocity data were used to detect individual blade tip vortices with ages above 450° close to the ground and to extract vortex parameters. For a takeoff maneuver, both concentrated vortices and the formation of larger vortex structures due to bundling of several vortices were observed.

Abstract An investigation into blade tip vortices of a sub-scale rotor under cyclic pitch conditions is carried out. Background oriented schlieren (BOS), particle image velocimetry (PIV), and computational fluid dynamics (CFD) are applied to the same test cases. This approach allows to combine the velocity data from PIV in a measurement plane, the density related data from BOS in a measurement volume, and the comprehensive set of flow variables provided by unsteady detached eddy simulations. Vortices up to an age of $$\varPsi _{{\text {v}}}= {70}^\circ$$ Ψ v = 70 ∘ in case of PIV and CFD, and up to $$\varPsi _{{\text {v}}}= {200}^\circ$$ Ψ v = 200 ∘ in case of BOS are considered. The vortex locations are obtained through all three techniques. The unsteadiness of the vortices was obtained by the experimental results, whereas CFD provides an average solution. An increased position scatter was observed during the downstroke of the pitch cycle with both experimental methods and was found to be in good agreement. In the second part, the PIV velocity data are compared to common vortex models. An approach to link the density distribution and the swirl velocity is applied to the measured data. Based on the CFD results, it is shown that the assumption of isothermal flow yields better agreement between velocity and density than isentropic flow. Graphic abstract

Measurements of the flow field around a free-flying model helicopter in ground effect for both quasi-steady and unsteady maneuvering flights were performed using stereoscopic particle image velocimetry (PIV). The wake features for hover and forward flight at low advance ratios were characterized and changing flow patterns like recirculation and ground vortex flow were observed to be in good agreement with existing wind tunnel data. Parameters describing both general flow patterns and single blade tip vortices were extracted and showed significant dependence on the forward flight velocity. Landing approaches were performed and large vortical structures were observed close to the rotor disk, which contained high amounts of vorticity due to entrained blade tip vortices. The vortex structures developing for unsteady landing approaches contained distinctly higher velocities and momentum fluxes than expected from quasi-steady conditions at the same advance ratios. For a vertical takeoff maneuver with a rapid increase of collective pitch, the bundling of blade tip vortices into a "starting vortex" with circulation values up to 6 times higher than for a single blade tip vortex was observed. These results show the significance of experimental data on free-flying helicopters in unsteady maneuvering flight because the resulting flow fields cannot be predicted using steady-state experiments or simulations.

A well-defined reference set of data for computational fluid dynamics and comprehensive code validation for a scaled helicopter main rotor with boundary layer transition in forward flight is presented. The boundary layer transition was measured using differential infrared thermography (DIT) on the top (suction) side of the NASA/Army “PSP rotor” in the NASA Langley 14-by-22-Foot Subsonic Tunnel. The tests used a FLIR X8500 SLS long-wave infrared camera to observe the three-bladed rotor. The boundary layer transition was detected for forward flight at an advance ratio of 0.3 (115 kt). The measured boundary layer transition positions are consistent with previous measurements and predicted boundary layer transition locations. A method for the analysis of DIT images for a rotor in forward flight is shown and validated based on computational analysis of a pitching airfoil with varying inflow, showing both qualitative and quantitative similarity to experimental data.

To characterize the vortex system of a helicopter rotor with cyclic pitch, an experimental investigation is carried out at the rotor test facility of the DLR, German Aerospace Center in Gottingen. ...

A full-scale wind-tunnel test of a replica of Otto Lilienthal’s Normalsegelapparat from 1893 was conducted in a closed low-speed wind tunnel. The goal was to determine the flight performance and ch...

In the published version, graphical abstract was missing. It is given below.

Abstract A method of measuring density second derivative (shadowgraph) as a variant of background oriented schlieren (BOS) is presented. Pairs of images of a moving background with a small time difference are analysed using moving-window cross-correlation, giving the second derivative of density in the axis of background travel. The advantage over standard BOS is that no reference image is required. In contrast to the existing “reference-free BOS” method, no double-image is produced, making the data interpretation simpler. The sum-of-correlation (ensemble correlation) method can be used on an image series, where the flow is constant over the series time, to significantly reduce the size of the cross-correlation window. The reduction in the window size and sharpening of the resulting image is similar to that achieved by optical flow, but this method is easily automatable, without knowledge of the flow, and can be used with any existing BOS analysis software. Graphic abstract

RAFFEL and Merz [1] demonstrated the feasibility of unsteady boundary-layer transition detection by differential infrared thermography (DIT), a method whereby the difference of two sequentially recorded infrared images is analyzed to extract the point of greatest difference between the flows, which, for attached flows, is equivalent to the transition position. For flows with dynamic flow separation, the strongest feature in the infrared difference images is no longer the transition, but the separation of the flow. Gardner and Richter [2] showed that the standard deviation of pressure measurements at the transition point has a small peak, but that separated flow results in an even larger peak. Similarly, the infrared images can be analyzed to extract the presence and extent of separated flowover an airfoil. DIT offers a great advantage over using pressure sensors as indicators for separated flow because DIT does not require any special equipment to be attached to or built into the model observed. The major aim of the investigations presented here is the generation of “stall maps” in which stalled areas in the rotor plane are geometrically described. Currently, these maps are often produced through the investigation of pressure sensor data from sensors integrated into the rotor blades [3]. Alternatives are the use of tuft visualization [4] or pressure-sensitive paint [5]. However, these techniques require installations or coatings on the rotor blades, which can be avoided by the technique described in the following

An analytical approach for the correction of the influence of permeability on aerodynamic lift and drag polars of finite wings is presented. The approach is validated through wind tunnel measurements of a simple wing-body configuration. Polars for both permeable and impermeable wings were measured at varying inflow velocities. Using only measured values for permeability and porosity, the permeable polars can be mapped to the impermeable case and vice versa with good agreement.

The presented work tackles the lack of experimental investigations of unsteady laminar-turbulent boundary-layer transition on rotor blades at cyclic pitch actuation, which are important for accurate performance predictions of helicopters in forward flight. Unsteady transition positions were measured on the blade suction side of a four-bladed subscale rotor by means of non-intrusive differential infrared thermography (DIT). Experiments were conducted at different rotation rates corresponding to Mach and Reynolds numbers at 75% rotor radius of up to$${M}_{75} = 0.21$$M75=0.21and$${\mathrm {Re}}_{75} = 3.3 \times 10^5$$Re75=3.3×105and with varying cyclic blade pitch settings. The setup allowed transition to be measured across the outer 54% of the rotor radius. For comparison, transition was also measured using conventional infrared thermography for steady cases with collective pitch settings only. The study is complemented by numerical simulations including boundary-layer transition modeling based on semi-empirical criteria. DIT results reveal the upstream and downstream motion of boundary-layer transition during upstroke and downstroke, a reasonable comparison to experimental results obtained using the already established$$\sigma c_p$$σcpmethod, and noticeable agreement with numerical simulations. The result is the first systematic study of unsteady boundary-layer transition on a rotor suction side by means of DIT including a comparison to numerical computations.

Boundary layer transition measurement was demonstrated using differential infrared thermography (DIT) on the top side of a helicopter rotor in forward flight, which detects the difference in the convective heat transfer at the boundary layer transition position. The tests used a FLIR X8500xc SLS long wave infrared camera to observe the DLR EC135 test helicopter rotor. The boundary layer transition was detected for hover out of ground effect (150 ft) and for forward flight at 80 kt (1700 ft). The measured boundary layer transition positions are consistent with previous measurements of the EC135 hovering in ground effect, and with predicted boundary layer transition positions. A method for the analysis of DIT images for a rotor in forward flight is shown, based on computational analysis of a pitching airfoil with varying inflow.

Helicopter operation close to the ground is of interest to the rotorcraft community as the rotor wake can lift up sediment particles from the ground, causing deteriorated visual conditions for the pilot - a phenomenon that is referred to as brownout. The impact of unsteady maneuvering flight on brownout is still unclear and quantitative flow field measurements are needed to better understand the underlying physical effects. The present work investigates particle image velocimetry (PIV) and the background-oriented Schlieren (BOS) technique for studying the rotor wakes of free-flying helicopters close to the ground. Two flight tests with a model helicopter and a full-scale BO 105 helicopter were conducted and time-resolved stereoscopic PIV and BOS data were acquired during steady and unsteady flight maneuvers in ground effect. This paper discusses the measurement setups realized in these flight tests, technical challenges of implementing the measurement methods in a flight test environment, and example flow field results. The presented measurement data and evaluated results demonstrate the successful first application of PIV, time-resolved BOS, and combination of both measurement techniques on free-flying model- and full-scale helicopters. This novel application of full-scale flow measurements enables a more complete investigation of flow phenomena in the wake of free-flying helicopters close to the ground and will contribute to better understanding rotor wake interactions such as brownout.

A rotor test facility was developed to investigate dynamic stall under optimized boundary conditions compared to conventional hover chambers. lt features a defined axial inflow, reduction of ground effect and recirculation of the rotor wake and good optical access to apply non-intrusive measuring techniques. The rotor consists of two blades with an aspect ratio of 6.8 and a tip radius of 0.65m. lt was operated at a Mach number of 0.21 and a chord based Reynolds number of 350000, both at 75% radius. The flow and blade deformation were analyzed by means of unsteady blade pressure transducers, particle image velocimetry and tip deflection measurements covering the whole azimuth and different radii. Three measurement approaches to detect flow separation: tufts, differential infrared thermography and surface pressure analysis showed comparable results and flow separation over one half of the cycle. A radiusdependent separation point in time and an altering of the dynamic stall behavior near the tip were observed. The recirculation area remained close to the blade's surface with radial jet flow increasing towards the tip. A maximum increase of the effective angle of attack of 1.4° was observed due to blade deformation. For a fully attached flow case the laminar-turbulent boundary layer transition differed along the radius and moved over 70% of the blade's chord.

Laminar-turbulent boundary-layer transition is investigated on the suction side of Mach-scaled helicopter rotor blades in climb and analyzed in view of the effect of rotational forces. Transition p...

The tip vortex of a finite rotor blade-tip model undergoing pitch oscillations is investigated by means of a wind-tunnel study. The motion and flow parameters as well as the model geometry were adapted to rotorcraft applications, including test cases with fully attached flow, light dynamic stall, and deep dynamic stall. High-speed particle image velocimetry was used to study the spatiotemporal behavior of the tip wake at different streamwise positions. Combined with surface pressure measurements, the data establish a connection between the sectional lift of the model and the vortex parameters (for example, swirl velocity and circulation), including dynamic phenomena like hysteresis and vortex breakdown during stalled flow conditions. For pitch oscillations below static stall limit, only small hysteresis effects were observed, and the vortex structure is similar to static reference measurements. Beyond static stall, the tip vortex retains its general structure in a time-averaged or phase-averaged frame, bu...

This section discusses various applications and techniques used in \(\mu \text {PIV}\). The applications include very high spatial resolution measurements of pressure-driven flow in a rectangular capillary and measurements in a challenging toroidal vortex.

The following PIV application performed in 1996 in a turbulent boundary layer at the wall of a flat plate illustrates two problems: obtaining PIV data close to a wall and recovering PIV data even in a flow with gradients (due to the velocity profile of the boundary layer). In the present series of experiments the measurement position was 2.3 m downstream of a tripping region in the low-turbulence wind tunnel at the DLR-Gottingen research center [17]. At this position the turbulent boundary layer thickness δ was of the order of 5 cm, of which the lower 3 cm was imaged. At free stream velocities of 10.3, 14.9 and 19.8 m/s between 90 and 100 PIV image pairs were recorded. The evaluation of the recordings were made without windows shifting and window deformation which were not available at that time.

In the research field of aeroacoustics, fluctuations in turbulent flows which are part of a sound generating process can be identified by means of the so-called causality correlation technique [5,6,29]. In this approach, the coefficient matrix resulting from the calculated correlation coefficients between the acoustic pressure fluctuations in the far field and velocity fluctuations in a turbulent flow field is analysed. The temporal resolution of the coefficient matrix is set by the sampling rate of the pressure signal in the far field. The spatial resolution is determined by the spatial sampling rate of the measured flow quantity (Table 11.22).