c
TU Delft, The Netherlands, 2006
PREMATURE DECAY OF WAKE VORTICES WITH
DIFFERENTIAL- AND OSCILLATING FLAP SETTING
Guido Voß∗, Stefan Melber-Wilkending∗ and Ralf Rudnik∗
∗Deutsches Zentrum f¨ur Luft- und Raumfahrt Institut f¨ur Aerodynamik und Str¨omungstechnik Lilienthalplatz 7, 38108 Braunschweig, Germany
E-mail: Guido.Voss@dlr.de, Stefan.Melber@dlr.de, Ralf.Rudnik@dlr.de Web Page: http://www.dlr.de/AS
Key words: Wake vortices, TAU-Code, counter-/co-rotating, Crow-Instability
Abstract. This paper gives an overview of the research on wake vortex decay using Dif-ferential Flap Setting (DFS) and Oscillating Flap Setting (OFS). The concept of DFS deals with the small deflection of flaps and elevator to produce additional vortices which interacts with the vortices from wing, elevator, flaps and wing-body-junction. Numerical calculations with the A340-300 model at this topic are presented in the first part of the paper. On the other hand the OFS concept works with the periodic movement of DLC (direct lift control) flaps and/or ailerons to trigger inherent instabilities of the vortices. The basic mechanism of the introduction of the periodic disturbance into the flow field is a periodic alternation of the spanwise lift distribution, i.e. of the circulation and spacing ra-tios of the distinct vortex pairs in the extended wake vortex nearfield. This can be achieved both by pitch oscillation and oscillating DLC flaps. Pitch oscillation alone are sufficient for accelerating Crow instability as proven already in the 1970s with small aircraft. One can distinguish between long-wavelength and short-wavelength instabilities. The second part of this paper is truncated on short-wavelength instability examinations i.e. the Crow instability.
1 Introduction
World airports face a long-term capacity problem due to the air traffic control separa-tion regulasepara-tions presently in use. These air traffic control separasepara-tion regulasepara-tions determine the minimum longitudinal distance between two aircraft on the same flight path depen-dent on the maximum take-off weights of both aircraft in order to prevent the following aircraft from encountering potentially hazardous wake turbulence. The attempts to al-leviate wake vortices in the early seventies lost importance upon proving the system of empirically-found separation standards to be a suitable measure to ensure safe air traffic operations. Meanwhile, due to the increasing number of congested airports and the devel-opment of very high capacity transport aircraft, the research efforts on the wake vortex problem have risen again.
2 Differential Flap Setting - Numerical Investigations and Wind tunnel Ex-periment on A340-300
The following part is a synthesis of the near-field CFD results of DLR within the Euro-pean project AWIATOR [1] and shows numerical results of the solution of the Reynolds-averaged Navier-Stokes (RANS) equations for the baseline high-lift and a differential flap high-lift configuration of an AIRBUS A340-300 ([2], [3]). The results from the RANS simulations served as input-data for rollup-studies. Besides the grid generation and the flow solution method the comparison between numerical results and measurements and a comparison between both configurations regarding the vortex wake are shown. Further on, the effects of different free stream conditions (Reynolds number effect) are shown in this report.
2.1 Configurations
The model used is the AIRBUS A340-300 1:27 catapult model of ONERA catapult tunnel, a wing span of b/2 = 1.118 m and an aerodynamic mean chord of c = 0.269 m. The free stream values of the computations are those of used in the catapult tests: V = 25 m/s, Re = 520 000. With the same configuration and a scale of 1:1 the real aircraft flight conditions are simulated with V = 65.8 m/s and Re = 36.94 · 106. The angle of attack is
adjusted iteratively during the flow solution process to achieve the same lift (CL = 1.4)
for all considered configurations, table 1.
2.2 Numerical Method
The RANS-equations are solved by the DLR hybrid unstructured flow solver TAU, which is based on a three-dimensional finite volume scheme. The governing equations are solved on a dual background grid, which, together with the edge-based data structure, allows to run the code on any type of cells. The solver is part of the MEGAFLOW-software package [5]. The temporal gradients are discretized using a multi-step Runge-Kutta scheme. The inviscid fluxes are calculated employing a central method with scalar dissipation. The viscous fluxes are discretized using central differences. In order to accel-erate the convergence to steady state, local-time stepping and a multigrid technique based on agglomeration of the dual-grid volumes are employed. Because of the low free-stream Mach-number (M = 0.076) preconditioning was used to improve accuracy and to enhance the convergence to steady state [6]. The target lift coefficient for both configurations is CL = 1.4, which was obtained by changing the angle of attack during the simulation
with the ”target lift” option of the TAU-code. The turbulence closure for the solution of the RANS-equations for the AIRBUS A340-300 is computed by using the one equation Spalart-Allmaras turbulence model [7] with Edwards modification [8]. To improve the quality of the vortex simulation a so called vortex correction for the Spalart-Allmaras turbulence model was used, [9].
The grids for the numerical simulations were generated with the hybrid unstructured grid generator package Centaur from CentaurSoft [10]. The grids consist of two parts: a quasi-structured prismatic cell layer in order to achieve an appropriate resolution of the viscous effects inside the boundary layer. In contrast to this, tetrahedral cells fill the outer domain of the flow-field.
To improve the resolution of the vortex wake six successive grid refinements by local adaptation are accomplished. The pressure loss was used as adaptation sensor, which gives the best results for vortex adaptations. Because this sensor would place the most of the new points in the boundary layer where vorticity is produced, a modification was implemented to mask out this areas of the grid from the adaptation procedure.
2.3 Numerical Results
In this section the results of the numerical simulations of both A340-300 configurations are shown. Besides the differences in the results for both configurations, the comparison with experimental results and the effect of different free stream conditions (Reynolds number effect) are shown and discussed.
The different flap setting of the baseline (case 1, compare table 1) and the DFS configu-ration (case 4) leads to a changed vorticity distribution, depicted in figure 3 at x = 2.65 m. The different positions and strengths of the vortices for both configurations can be clearly seen in this figure. The main difference between both configurations is a merged flap and nacelle outboard vortice in the baseline configuration compared to individual vortices in case of the DFS configuration.
The next comparison shows the effect of catapult (case 2) and flight Reynolds (case 4) number on the vortex topology and the lift distribution for the differential-flap configura-tion. To reach the same lift coefficient CL = 1.4 at catapult condition an angle of attack
of α = 10.94o and at free flight condition an angle of attack of α = 7.13o is needed. The
vortex topology (figure 4) is different for both cases: the wake behind the configuration is more pronounced at flight Reynolds number, the vortices are not so concentrated in this case. Comparing both figures an increased noise is obvious for the flight Reynolds number case in the flow field around the wake. This might be an effect of the increased Reynolds number with its decreasing viscosity, resulting in a decreased natural dumping of vortex fragments in the flow field. A second effect can be the more concentrated vortex sheet with its higher gradients in case 4, which leads also to an increased noise.
3 Differential Flap Setting and Oscillating Direct Lift Control - ATTAS flight test
Within the framework of the European research project AWIATOR a test with the Sanders smoke pod installed on the DLR-ATTAS (Advanced Technologies Testing Aircraft System) was carried out to investigate the effect of differential flap setting (DFS) and oscillating direct lift control (DLC) flaps on respectively the vortex wake visualised by the smoke trace.
The research aircraft DLR-ATTAS is a modified short haul passenger jet VFW 614. The aircraft was converted into an in-flight simulator and technology demonstrator by DLR and MBB in 1981-1986. The aircraft span is 21.5 m, maximum take-off weight is 20.95 t. The high lift system of the DLR-ATTAS consists of a main landing flap and integrated six independent DLC flaps. Main flap operation specifications are:
- possible flap settings δf lap: IN, SP, 1o, 5o, 14o, 35o
- vmin at δf lap = 14o: vmin = 1.2 ∗ vs≈ 98 kts ≈ 50m/s
The DLC flaps are pairwise independently movable flap tabs. They allow for the real-ization of static and oscillating differential flap setting. The DLC operation specifications are:
- for DLC operation the main landing flaps must be deflected by SP ≤ δf lap ≤ 14o
- DLC operation must be symmetric on both wings
- DLC max deflection range is ±36o, at approach speed is ±15o
- DLC deflection can be selected in steps of ∆δDLC = 2.5o
The DLC flaps are numbered from left to right, i.e. the outboard DLC on the left wing is No. 1, the outboard DLC on the right wing is No. 6.
3.1 Numerical Calculations
Prior to the ATTAS smoke test numerical investigations were conducted in order to allow a reasonable selection of configurations to be in flight test. In a first step the spanwise lift distributions of the ATTAS for several configurations are calculated by DLR using the vortex lattice code Lifting-Line. Secondly, the roll-up of the vorticity distributions corresponding to the determined lift distributions is calculated with a vortex method by the Universite de Louvain (UCL). The evaluated circulation and spacing ratios of the rolled-up vortex wake in two vertical planes at x/b = 3.5 and x/b = 23 serves as input for instability calculations performed by ONERA.
3.2 ATTAS Flight test configurations
From the preceding numerical investigations five configurations are selected to be tested in flight. An overview of all tested configurations is given in figure 5 and 6. The reference case (Configuration 1) is the baseline ATTAS configuration in approach (flap deflection δf lap = 14o and DLC deflection δDLC = 0o).
In static DLC mode are tested a co-rotating/inboard loading (Configuration 2) a counter-rotating/outboard loading (Configuration 3). In oscillating DLC mode two con-figurations are considered with the low frequency corresponding to the Crow wave length are flight: First the co-rotating/inboard loading (Configuration ) and second the counter-rotating/outboard loading (Configuration 5).
3.3 Programming and simulator test of DLC functions
To cover the ATTAS Flight test configurations different parameters and functions for the control of the DLC deflections have been defined. The respective mathematical formu-lations were coded in FORTRAN and validated using the ATTAS ground simulator. The selection of DLC behaviour can be triggered by predefined switches (”ERR-Switches”).
In case of the inboard loading configuration (Conf. 4) an upward (negative) pre-deflection of the DLC flaps pair 1/6 can be selected in a range between −37.5o ≤ η
DLC1/6 ≤
0o(physical deflection limitation of unloaded DLC flaps pair 1/6 is −36o ≤ ηDLC1/6 ≤ 35o).
The equivalent (positive) downward pre-deflection of the DLC flaps pairs 2/5 and 3/4 is calculated by
ηDLC2/5 = ηDLC3/4 = −
S1+ S2
S2+ S3+ S4+ S5
∗ ηDLC1/6
configuration of fixed DLC setting within the above defined range for the DLC flaps pair 1/6 can be selected in increments of ∆ηDLC1/6 = −2.5o. The corresponding correct value
for the DLC flaps pair 2/5 and DLC flaps pair 3/4 will be automatically established by the experimental software.
Starting from the fixed pre-deflection settings of the DLC flaps an oscillation can be initiated in a frequency range between 0.05 Hz ≤ ω ≤ 3.15 Hz in increments of ∆ω = 0.05 Hz. The actual amplitude A of the oscillation is in between the maximum amount of deflection Amax of the respective DLC flaps pair and zero deflection
In case of the outboard loading configuration (Conf. 5) the DLC flaps pair 1/6 is deflected downwards (positive) into a fixed position in a range between 0o ≤ η
DLC1/6 ≤
37.5o. The adjustment of respective lift variation has to be made by the pilot. The
(positive) pre-deflection of the DLC flaps pair 2/5 is identical with the DLC flaps pair 1/6ηDLC1/6 = ηDLC2/5. The equivalent (negative) upward pre-deflection of the DLC flaps
pair 3/4 to compensate for the lift variation produced by the DLC flaps pair 2/5 is calculated by
ηDLC3/4 = −
S2+ S5
S3+ S4
∗ ηDLC2/5
Again this approximation was used to keep the lift roughly constant when the DLC flaps pair 2/5 move against the DLC flaps pair 3/4. Any configuration of fixed DLC setting within the above defined range for the DLC flaps pair 1/6 can be selected in increments of ∆ηDLC1/6 = 2.5o. The corresponding correct value for the DLC flaps pair
3/4 will be automatically established by the experimental software.
Starting from the fixed pre-deflection settings of the DLC flaps an oscillation can be initiated for the DLC flaps pairs 2/5 and 3/4 while the DLC flaps pair 1/6 remains fixed. Again the frequency range is between 0.05 Hz ≤ ω ≤ 3.15 Hz and can be varied in increments of ∆ω = 0.05 Hz.
3.4 Flight procedure
• a line-up area which is part of the flight track of the ATTAS inbound to the waypoint A, this line-up area extends over 6.2 NM between waypoints D and E.
• a test area extending for 3.15 NM (about 5800 metres), from waypoint E to A. In this area the smoke will be generated and the length of this segment is equal to the length of the smoke plume
• a second line-up area meant for the NLR Citation, extending from the waypoint I to J, in this line-up area the airspeed of the NLR Citation will be decreased from 180 KIAS to 130 KIAS, such that the aircraft will fly straight and level at 130 KIAS at approximately waypoint J, in this line-up segment the Citation crew has to adjust the airspeed depending on the relative distance between the Citation and the oncoming ATTAS
Details of the flight procedure can be found in Wever [11]. 3.5 Data Recording
Besides the video recording done with a video system onboard of the NLR Citation the smoke trace development was documented from the ground with two photo cameras and a video camera. Pictures were taken with a digital camera Minolta MLT100 and for very high image resolution with a Rolleiflex 6008 AF. The video recording was done with a Panasonic NV MX300EG. The ground equipment was located parallel to the runway (distance approx. 150 m) and 200 m behind the threshold.
For the understanding of vortex dynamics the video motion pictures are helpful, how-ever, the quantitative analysis is easier done using photos which are taken in constant time intervals. The image resolution of the Minolta digital camera was found to be sufficient. Thus, in the following discussion of results mainly the Minolta camera photos are used. The observation from the ground was impeded on both days by clouds which blocked the view such that no complete documentation of the flight test could be obtained.
The analysis of the smoke visibility has been done from two viewpoints, on the one hand from the cabin of the ATTAS and the other hand from the ground. The judgement of the smoke visibility from the cabin was difficult due to the limited viewing angle. Therefore the analysis will be done using the images taken with the ground cameras.
3.6 Analysis of smoke trails
A comparison of both patterns is shown in figure 8 - figure 10 (0 sec - 90 sec, trig-gered Crow instability f = 0.5 Hz). While in the static case (left column) no upcoming instability is observed, in the oscillating case (right column) after 30 sec the triggered long-wave instability becomes visible. After 60 sec with oscillating DLC flaps distinct half rings are seen (smoke generator only on left wing) which indicates the formation of vortex rings. A complete ring can be identified at the very left side in Fig. 10 after 90 sec. In the undisturbed case the topology after 75 sec is comparable to the topology of the oscillating case after 30 sec. Thus, a substantial speed-up of the growth of instability is achieved.
4 Conclusion
This first part shows the numerical flow simulation of high lift flows for AIRBUS A340-300 configuration with the DLR TAU-code. The resolution of the vortex wake behind the aircraft was increased using grid adaptation. With this technique grid independent solutions can be obtained. Both considered configurations, the approach configuration and a differential flap configuration, were simulated at the same lift coefficient using the target lift technique of the TAU-code. The numerical results are compared to measurements in the DLR NWB low speed wind tunnel. Furthermore both configurations are compared against each other in respect of differences in the lift distribution and the flow field behind the aircraft. Finally the effect of different free stream conditions (Reynolds number effect) is shown an discussed.
In the second part a flight test with the DLR-ATTAS with installed smoke pod for visualisation of the vortex wake and the NLR-Citation in the framework of the European integrated project AWIATOR at Braunschweig airport is shown. The results of the NLR-Citation video recording is presented in a separate report, see Wever [12].
The effect of oscillating DLC flaps causing a periodic alternation of spanwise lift dis-tribution has been investigated. First, numerical studies have been performed to define promising configurations. After determination of spanwise lift distributions with a vortex lattice method the roll-up was calculated with a vortex method and finally the linear instability phase was determined for the rolled-up vortex topologies. From the numerical investigations ATTAS- configurations were derived for from the programming of the DLC flaps in the ATTAS flight control system. The proper DLC function was tested in the ATTAS ground simulator prior to the flight test.
easily separated. REFERENCES
[1] AWIATOR - Aircraft Wing With Advanced Technology Operation. www.awiator.net. [2] Melber-Wilkending, S.: AWIATOR: Technical report on Extraction of the local-Lift Coefficient from RANS-Simulations. AWIATOR correspondence, AW-DLR-111-001F.
[3] Melber-Wilkending, S.: RANS-Solutions of complex configurations as Input for Vortex-Rollup Studies III. AWIATOR Technical Report, AW-DLR-111-004F, 2004. [4] Stumpf, E.; Kuepper, A.; deGroot, K.: Equilibirum Test of the A340 ONERA
cata-pult model in DNW-NWB. AWIATOR correspondence, AW-DLR-112-001F.
[5] Kroll, N.; Rossow, C.-C.; Becker, K.; Thiele, F.: MEGAFLOW - A Numerical Flow Simulation System. 21st ICAS congress, 1998, Melbourne, 13.09-18.09.1998, ICAS-98-2.7.4, 1998.
[6] Melber, S.; Heinrich, R.: Implementierung einer Praekonditionierungstechnik f¨ur Str¨omungen kleiner Mach-Zahlen im DLR TAU-Code und Anwendung im Hochauftrieb. 13th AG STAB/DGLR Symposium M¨unchen, 12.-14. November 2002. [7] Spalart, P. R.; Allmaras, S.R.: A One-Equation Turbulence Model for aerodynamic
Flows. La Recherche Aerospatiale, Nr. 1, 5-21, 1994.
[8] Edwards, J.R.; Chandra, S.: Comparison of Eddy Viscosity-Transport Turbulence Models for Three-Dimensional, Shock-Separated Flowfields. AIAA Journal, No. 4, April 1996.
[9] Melber, S.: Wirbelkorrektur fuer Ein- und Zweigleichungs-Turbulenzmodelle und Im-plementation fuer das Spalart-Allmaras Turbulenzmodell in den Stroemungsloeser DLR-TAU. DLR IB 124-2002/17, 2002.
[10] Kallinderis, Y.: Hybrid Grids and Their Applications. Handbook of Grid Generation, CRC Press, Boca Raton / London / New York / Washington, D.C., pp. 25-1 - 25-18, 1999.
[11] Wever R., Krawal A. K.: Flight test plan for the NLR-DLR flight test program in preparation for the Flight Test I in the AWIATOR project. NLR-CR-2003-443 or AW-NLR-113-006, 2003.
5 Tables
case Case 1 Case 2 Case 4
configuration baseline differential-
differential-flap flap
mode Navier- Navier-
Navier-Stokes Stokes Stokes
angle of attack α 9.13o 10.94o 7.13o Velocity V [m/s] 25 25 65.8 Reynolds num. Re [-] 0.52 · 106 0.52 · 106 36.94 · 106 slat 23o 23o 23o flap inboard 26o 17o 17o flap outboard 26o 32o 32o aileron 5o 5o 5o elevator setting −5.75o −10.5o −10.5o
Table 1: Settings / free-stream values of A430 configurations.
6 Figures
Figure 2: Total pressure loss ∆ptotfor baseline configuration, DLR TAU-code (x = 2.9 m) and DLR-NWB
Figure 4: Cut-plane at x = 2.65 m (x/b = 0.84) perpendicular to the fuselage axis, vorticity ωx, Left:
Case 2, DFS configuration (17/32), RANS simulation, Re = 0.52 · 106, V = 25 m/s, α = 10.94o, Right:
Case 4, DFS configuration (17/32), RANS simulation, Re = 36.94 · 106, V = 65.8 m/s, α = 7.13o.
Figure 6: Oscillating configuration
Figure 9: Comparison static vs oscillating DLC 45...60s
