ABSTRACT NOMENCLATURE 1.0 INTRODUCTION THE AERONAUTICAL JOURNAL FEBRUARY

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1 THE AERONAUTICAL JOURNAL FEBRUARY Experimental investigation of the effect of nozzle shape and test section perforation on the stationary and non-stationary characteristics of flow field in the large transonic TsAGI T-128 Wind tunnel V. I. Biryukov, S. A. Glazkov, A. R. Gorbushin, A. I. Ivanov and A. V. Semenov Central Aerohydrodynamic Institute, TsAGI, Zhukovsky, Russia ABSTRACT The results are presented for a cycle of experimental investigations of flow field characteristics (static pressure distribution, static pressure fluctuations, upwash, boundary-layer parameters) in the perforated test section of the transonic TsAGI T-128 Wind Tunnel. The investigations concern the effect of nozzle shape, wall open-area ratio, Mach and Reynolds numbers on the above-outlined flow characteristics. During the tests, the main Wind-tunnel drive power is measured. Optimal parameters of the nozzle shape and test section perforation are obtained to minimise acoustic perturbations in the test section and their non-uniformity in frequency, static pressure field non-uniformity, nozzle and test section drag and, accordingly, required main Wind-tunnel drive power. NOMENCLATURE A c C p d f h M P rms St amplitude sonic speed pressure coefficient perforation orifice diameter open-area ratio; also, frequency test section height Mach number pressure root mean square Strouhal number U, u longitudinal flow velocity component V, v vertical flow velocity component W Wind-tunnel main drive power X,Y,Z longitudinal, vertical and transversal coordinates α angle-of-attack; pitching angle; upwash β slip angle, sidewash γ rolling angle δ boundary-layer thickness δ * boundary-layer displacement thickness δ ** boundary-layer momentum thickness θ test section wall angle Indices e conditions at boundary-layer outer edge t parameter at stagnation point conditions in free-stream flow 1.0 INTRODUCTION The TsAGI T-128 Wind tunnel is intended to test models of vehicles and their components at Mach numbers from 0 15 to 1 7 in a wide range of Reynolds numbers. Figure 1 shows the general arrangement of the Wind tunnel. The T-128 Wind tunnel is a continuous-operating close-circuit facility. Its schematic is A version of this paper was presented at the 2004 RAeS Aerospace Aerodynamics Research Conference.

2 76 THE AERONAUTICAL JOURNAL FEBRUARY Cooler 2. Inlet and ou tlet 3. Compressor 4. Deturbulising screens 5. Variable no zzle Figure 1. T-128 Wind-tunnel. 6. Test section 7. Test section flaps 8. Turning blades 9. Honeycomb 10. Settling chamber 11. Hatch for test section change 12. Plenum chamber suction 13 Suction outlet 14. Diffuser 15. Protecting screens Figure 2. T-128 Wind-tunnel scheme. Figure 3. Total pressure vs Mach number envelope. presented in Fig. 2. The basic Wind-tunnel components are as follows: a main drive with power of 100Mw, a four-stage compressor, an air cooler, turning vanes, a honeycomb, a plenum chamber, a variable nozzle, removable test sections. To extend the experimental potentialities, the systems of pressurising, vacuuming and forced gas suction from the plenum chamber are provided in the Wind tunnel. At present, four of five 2 75m by 2 75m test sections designed for the T-128 Wind tunnel are in service. They offer a wide spectrum of experimental capabilities of the facility. The unique design feature of the T-128 Wind tunnel is its capability of changing the porosity of the perforated walls of test sections No 1, No 2 and No 5 (test section No 3 has slotted walls of variable porosity). About 130 panels are provided on the perforated test section walls and their open-area ratio can be set individually. Since the first run of the Wind tunnel in December 1982, methodical investigations of the flow fields in the test section were carried out many times (1,2) and optimal geometric parameters of the facility components were determined for the purpose of further reduction in the test error level. For example, it is shown in (Ref 3) that at high subsonic velocities the minimum effect of the walls of T-128 test section No 1 on the aerodynamic characteristics of a passenger aircraft model is noted at f = 3% to 6%. In particular, the integral upwash on the model wing surface due to the perforated boundaries proves to be equal to 0 at f = 3%, while zero upwash gradient over the span of the typical passenger aircraft model wing is obtained at f = 6%. The great range of the investigations is motivated by the diversity of test types and the great number of variable Windtunnel parameters. In particular, variable parameters are side wall angles in test sections No 1 and No 2 and position of test section flaps, configuration of flexible side nozzle walls, intensity of gas suction from the boundary-layer on the side walls ahead of the 2D-model in test section No 3 or from the plenum chamber in other test sections etc. The next cycle of investigations of flow characteristics in test section No 1 intended for testing complete models on the rear and fin stings was undertaken in October and November The investigations involved the effect of nozzle shape, wall open-area ratio, Mach and Reynolds numbers on flow parameters (static pressure fluctuations, static pressure distribution, upwash, boundary-layer parameters on the nozzle and test section walls). During the tests, the Mach number was varied from 0 2 to 1 1. The Reynolds number was varied by changing stagnation pressure within the range of P t = 100 to 400kPa. The T-128 Wind-tunnel test condition envelope in terms of Mach number and total pressure (without application of the suction system) is given in Fig. 3. The markers indicate conditions for the given test series. During the tests, the wall open-area ratio was varied for the whole test section from 1 to 18%.The pressure distributions were measured on the test section and nozzle walls, as well as on a short static pressure probe placed along the Wind-tunnel axis in the central unit of the α-mechanism. The probe has a five-point hemispherical head to measure upwash and sidewash. The total pressure profiles in the boundary-layer were measured in two sections: at the test section entry and on the side wall at the model locations. In addition to average flow characteristics, the unsteady pressure components in the band up to 5kHz were measured at several characteristic points on the test section walls, in the nozzle and in the plenum chamber, using Kulite transducers. The T-128 nozzle is designed with flexible side walls. The upper and lower nozzle walls are parallel. At sub- and transonic velocities, the tests were conducted for standard and extended nozzle positions (in the latter case, the nozzle contraction region was displaced towards the settling chamber by 4 to 5m). Both configurations of the nozzle and the test section are presented schematically in Fig. 4. The arrangement of unsteady pressure gauges, No 1 (on the side nozzle wall), No 2 (in the region of the reference plane of the test section), No 3 (in the test section mixing chamber flaps), No 4 (in the plenum chamber at some distance from the perforated

3 IVANOV ET AL EXPERIMENTAL INVESTIGATION OF THE EFFECT OF NOZZLE SHAPE & TEST SECTION PERFORATION test section wall) and No 5 (on the plenum chamber wall) and the rakes to measure boundary-layer parameters are also shown. Figure 5 illustrates the photo of the test section with a short probe on the axis and the nozzle at the extended position. The arrows indicate the arrangement of rakes to measure the boundarylayer profile in the nozzle and in the test section, as well as the static pressure fluctuation gauges No 1 and No 2. A diagram of the variable perforation method is shown top right. Simultaneously with investigation of the flow fields, the main drive power was measured for different conditions. This made it possible to evaluate power consumption for each condition and to identify configurations optimal in terms of power saving. Figure 4. Nozzles and test section configuration. 2.0 MEASUREMENTS OF STATIC PRESSURE DISTRIBUTIONS In test section No 1, the static pressure was measured on all the walls along the central generating lines, as well as on the upper and lower walls of the nozzle. In addition, a short probe to measure static pressure was placed additionally on the Wind-tunnel axis at the model location. The pressure distributions obtained allow the calculated methods (3) intended to determine wall interference and effect of supporting devices to be verified. The system of supporting devices is simulated by distributed singularities the intensity of which depends on geometry of the strut and supporting device. The perforated test section walls also have singularities and their intensity is derived from the Darcy-type boundary condition for a permeable surface. Compressibility is taken into account in the form of the Prandtl-Glauert rule. For example, Fig. 6 compares the distributions of pressure coefficients C p measured by the probe and obtained by numerical calculation of flow for two velocity conditions M = 0 7 and M = 0 9 at f = 6%. The calculation simulates compressible flow, limited by perforated walls, over the central sting unit (with probe) and pitch sector in which the sting is mounted. The calculated and experimental data are in adequate agreement, which indirectly confirms the correctness of the calculated evaluation of the perforated walls effect. The measurements make it possible to perform comparative analysis of the effect of the perforation ratio coefficient f on flow uniformity at the model location (co-ordinate of model rotation axis X = 4 75m). Figure 7 demonstrates the distributions of pressure coefficient C p, measured on the lower wall and using the probe, as a function of the wall porosity characterised by the open-area ratio. The corrections excluding only interference of the supporting devices (without consideration of the wall effect) are applied to the measurement results. The data for M = 0 85 illustrate the improved flow quality at f = 6% as compared with f = 2% and f = 10%. Rms deviations of C p from the mean average value and that measured on the probe 1. Short axial probe 2. Boundary layer rake on side wall 3. Unsteady pressure gauges 4. Boundary layer rakes in the nozzle 5. Variable perforation diagram Figure 5. Hardware installation. rms C = p 1 k= n C n 1 k = 1 2 p are given in the table below. f, % rms C p The results for the nozzle shape effect on the flow uniformity at the beginning of the test section are presented on the plots of Fig. 8. These plots show the Mach number distributions on the lower wall calculated using the static pressure distribution. The extended nozzle provides a smoother flow transition from the nozzle to the test section and approximately halves the amplitude of deviations from a specified Mach number value. Figure 6. Probe pressure distribution.

4 78 THE AERONAUTICAL JOURNAL FEBRUARY 2005 Figure 10. rms C p in test section (gauge No 2) Figure 7. Pressure distributions along the walls and probe. Figure 8. Mach number distribution on the bottom wall. Figure 11. rms C p (M = 0 85). Figure 9. Nozzle rms C p. Figure 12. Fluctuation spectra, obtained at different open area ratios.

5 IVANOV ET AL EXPERIMENTAL INVESTIGATION OF THE EFFECT OF NOZZLE SHAPE & TEST SECTION PERFORATION MEASUREMENTS OF STATIC PRESSURE FLUCTUATIONS Dependence of rms static pressure fluctuation deviations on the Mach number in the nozzle (transducer No 1) for atmospheric stagnation pressure is presented in Fig. 9 for different open-area ratios. Changes in the open-area ratio within the range of small values (f = 1% to 6%) do not exert much effect on the rms pressure coefficient deviations. The maximum fluctuation level for these open-area ratio values is noted at M = 0 6 (rms C p ~ 1 2% to 1 3%). As the free-stream flow velocity rises, the perturbations reduce drastically in the nozzle down to rms C p ~ 0 25% at M 1. There is also a local maximum at M = 0 3. At small Mach numbers, an increased fluctuation level is exhibited at f = 18%, while at transonic velocities there are local maxima for f = 10% at M = 0 9 and for f = 18% at M = Similar dependencies for transducer No 2, used to measure pressure at the beginning of the test section, are given in Fig. 10. At small wall open-area ratios and small flow velocities, the pressure fluctuation level in the nozzle and in the test section is almost the same. As the Mach number rises, the background noise in the test section increases to attain a maximum value rms C p = 1 8% at M = 0 8 to With further Mach number increase, the fluctuation level reduces. A drastic decrease in the perturbations in the nozzle at transonic flow velocities is explained by the fact that the main noise source is the test section and, as the Mach number rises, penetration of acoustic noise to the nozzle from the test section decreases. Variations in the open-area ratio from 1% to 6% does not have a significant effect on the fluctuation level in the test section over the whole range of the Mach numbers, as in the nozzle. An increase in the open-area ratio to f = 10% or 18% results in a sharp rise in the noise level in the test section. At high subsonic flow velocities, there exist local maxima of rms pressure coefficient deviations in the test section at f = 10% and f = 18% for the same Mach numbers as in the nozzle. Figure 11 shows pressure coefficient fluctuations as a function of the open-area ratio for all the transducers at M = The noise level in the nozzle (transducer No 1) and in the plenum chamber (transducers No 4 and No 5) is almost the same over the whole range of wall open-area ratios. The maximum perturbation level is noted in the zone of the mixing chamber flaps (transducer No 3), where a layer of flow mixing with gas in rest in the plenum chamber forms. As the open-area ratio rises, the perturbation level increases drastically in the test section. The minimum perturbation level in the nozzle (transducer No 1) and in the test section (transducers No 2 and No 3) is measured at f = 6%. The tests at increased stagnation pressures did not reveal any significant stagnation pressure effect on the pressure coefficient fluctuation level in the nozzle and in the test section. Figure 12 illustrates typical acoustic noise spectra in the test section (transducer No 2) at M = 0 85 and several open-area ratios. At small open-area ratios, there exists a fundamental tone at f ~ 2,700Hz and two small peaks with frequencies of f ~ 3,700Hz and f ~ 4,900Hz. With an increase in the open-area ratio, the perturbation maximum shifts towards the region of f ~ 90Hz and 150Hz with a drastic growth of perturbations in the intermediate region.the region of acoustic perturbations in the test section can conditionally be divided into two parts: the region of low frequencies (f < 1kHz) and the region of high frequencies (f > 1kHz). In the low-frequency region, the perturbation source is a fourstage compressor, which causes a wide spectrum of perturbations because of a great number of blades. The compressor-caused acoustic frequencies are derived from the well-known formula (4) : nm f = 60 where n is the compressor rotational speed (rpm), m is the number of blades. In Ref. 4, the well-known formula to obtain resonance frequencies in the test section is also presented. For the T-128 square test section, this formula is as follows: c f = h k l where k, l are integers. A peak is noted at f ~ 90Hz for f = 10% in the nozzle, test section and plenum chamber. The resonance tone in the test section k = l = 2 corresponds to this peak. In addition to the fundamental tone, there exists increased wideband noise within the band from 100 to 150Hz which corresponds to the tone in the test section k = 2, 3 and l = 1, 2, 3... As the wall open-area ratio rises up to f = 18%, there occurs a strong peak in the test section f ~ 130Hz, to which the tone in the test section k = l = 3 corresponds. Other tones in the test section are noticeable: f ~ 90Hz (k = l = 2), f ~ 170Hz (k = l = 4), f ~ 200Hz (k = l = 5), f ~ 260Hz (k = l = 6). For the open-area ratio of f = 4%, the fundamental tone of the compressor f ~ 290Hz is detectable, to which the tone of the test section k = 6, l = 8 and a small peak f ~ 150Hz (k = 5 and l = 3) correspond. The increased pressure coefficient fluctuation level at high wall open-area ratios is governed by a high acoustic perturbation level in the test section in the low-frequency spectrum region. In the high-frequency spectrum region, the tones are caused by interaction of the free-stream flow with sharp perforation orifice edges. The mechanism of formation of such tones is described in detail in the literature. Figure 13 shows the Strouhal number corresponding to highfrequency tones in the test section as a function of the Mach number. The Strouhal number is based on the perforation orifice diameter in the ordinary way: f d St = u The data are presented for f = 4% and 6%. For comparison, the data from Ref. 5 for the TWT ARA Wind tunnel are also given. This figure shows the well-known dependence of the Strouhal number for edge tones on the Mach number (5) : n St = for n = 1; 2; 3; 4; 5; 6....(1) 2 π(1+ M) In T-128 test section No 1, there occur four frequencies of acoustic vibrations in the upper part of the spectrum: f ~1,500Hz, 2,700Hz, 3,700Hz, 4,900Hz, the peak f ~ 2,700Hz (St ~ 0 2 to 0 3) having the greatest amplitude. In the region of small Mach numbers (M < 0 7), only the first two tones are excited, in the region of moderate Mach numbers (0 7 < M < 0 9) the second and the fourth tones are excited, while in the transonic region (M > 0 9) the second and third tones are excited, the third tone being excited only at an open-area ratio of f = 4%.The analysis of the data carried out for lower wall open-area ratios (f = 1% and 2%) has revealed that as a whole the situation corresponds to the case when f = 4%. Thus, from Figs 7 to 13, at transonic velocities, the perforation f = 6% proves to be optimal in terms of the number of excited tones and uniformity of acoustic perturbations in the test section in amplitude. Also, the open-area ratio f = 6% provides a minimum level of rms static pressure deviations in the test section. The Strouhal numbers given for TWT (f ~ 2 65kHz) are in agreement with the data obtained in the T-128 Wind tunnel. The dependencies calculated by Equation (1) have the same tendency as the experimental data but they are characterised by a weak dependence on the Mach number.

6 80 THE AERONAUTICAL JOURNAL FEBRUARY 2005 Figure 13 Strouhal number. Figure 16. Nozzle shape influence on the boundary-layer, M = 0 8. Figure 14. Boundary-layer displacement thickness at different wall inclination angles, M = 0 95, X = 3 86m, atmospheric total pressure. Figure 17. Upwash distribution in vertical plane, γ = 180. Figure 15. Velocity profiles in the boundary-layer on the right wall, X = 3 93m, M = Figure 18. Main drive power increment.

7 IVANOV ET AL EXPERIMENTAL INVESTIGATION OF THE EFFECT OF NOZZLE SHAPE & TEST SECTION PERFORATION MEASUREMENTS OF BOUNDARY- LAYER PARAMETERS In the present test series, the velocity profiles in the boundary-layer were measured at two sections. The first section was at the nozzle exit and had coordinate X = 0 8m relative to the beginning of the test section. In this section, the rakes were installed on the upper and lower walls. These rakes consisted of 22 Pitot probes to measure velocity profiles in the wall zone with a thickness of up to 140mm. One more rake was placed on the side wall at the model location (X = 3 93m) (Fig. 5). This rake was 280mm in span and comprised 18 total pressure probes. Relying on the measurement results, the profiles of Mach number and velocity U/U e in the boundary layer were constructed, as well as the integral parameters. The displacement thickness δ* and the momentum thickness δ** were calculated by the standard formulae: * U = δ ρ 1 dy ** U U δ and 0 eu = ρ δ ρ δ 1 dy e ρ 0 eu e U e The boundary layer on the walls of T-128 test section No 1 was investigated many times during methodical tests. As a whole, it is characterised by a thickness reaching δ ~ 130mm to 160mm for the standard configuration (subsonic nozzle, wall angle of 0 5º) at the model location. The boundary-layer has the velocity profiles generally typical for a rough surface. Displacement thickness values δ* at the model location are ~ 20mm to 25mm. In one of the previous investigations, one more parameter was varied, namely the side wall angle. Figure 14 presents the measured boundary-layer displacement thickness at section X = 3.86m as a function of the wall open-area ratio for four wall angles (θ = 0 5º, 0º, 0 5º and 2 0º). As for the standard wall installation providing zero pressure gradient at the model location (θ = 0 5º), a weak decrease in δ* is noted at f = 1 5% to 10% which is likely to be caused by a weak gas flow out of the test section. These data are obtained at M = 0 95, but the qualitative nature of the dependence also remains the same under other velocity conditions. The velocity profiles measured at the standard wall angle in the last test series also demonstrate the boundary layer growth for the maximum open-area ratio (Fig. 15). A change in the nozzle shape with the upstream contraction region displacement also results in boundary-layer growth. Figure 16 shows the velocity profiles measured at the nozzle exit and on the test section wall for the standard and extended nozzle positions. The difference in the displacement thickness obtained for the boundary layer at the nozzle exit remains the same at the model location, as well. 5.0 UPWASH MEASUREMENTS The conventional method of upwash measurements in the T-128 Wind tunnel implies model tests in the direct and overturned positions. In the present test series, upwash was measured using the static pressure probe installed in the central unit of the pitch sector. The head part of the probe was hemispherical and included five pressure taps: central tap to measure total pressure and two pairs of opposite taps drilled at 45º to the axis to measure upwash and sidewash. The five-point probe was calibrated during the tests. For this purpose, the probe was displaced in angle of attack within the range of ±2º in two positions: direct position (γ = 0º) and overturned position (γ = 180º). The resulting standard calibration characteristics enabled the data processing for all conditions under study. At zero pitch angle, the probe was along the test section axis and the fivepoint head directly measured upwash in the Wind tunnel. An example of upwash measurements on the Wind tunnel axis is given in Fig. 17 for different Mach numbers and different open-area ratios. Note that general upwash measurement results by testing the model in the direct and overturned positions are more conservative because they represent the parameter averaged over the major portion of flow. 6.0 MEASUREMENT OF WIND-TUNNEL MAIN DRIVE POWER During the tests, the Wind-tunnel main drive power was measured to determine optimal test section parameters providing minimum test section drag and, accordingly, facility-consumed power. It is shown in the Section devoted to pressure fluctuation measurements that the maximum acoustic perturbation level in the test section is obtained at M = Let us consider how the open-area ratio affects the main drive power at this Mach number. Figure 18 presents variations in the main drive power expressed in percent of maximum value as a function of open-area ratio. The power at f = 1% is taken as the zero value. The consumed power increases with the open-area ratio to attain maximum value at W ~ 5 6% at f = 18%. At moderate openarea ratios, there exists a local minimum W ~ 1 5% at f = 6%. As a whole, the curve shape is similar to the dependence of the rms static pressure fluctuation deviation in the test section on the open-area ratio given in Fig. 11. It means that the increase in the required main drive power is partially caused by formation of low-frequency acoustic perturbations in the test section at high open-area ratios. In addition, the perforated walls are a noise source and produce great drag, thus resulting in loss of flow mechanical energy when passing through the test section. The data presented on the boundary layer characteristics under different perforation conditions (Fig. 14) are in agreement with the measured power consumed by the main drive. Noteworthy is the correlation of the local minimum of the facilityconsumed power and rms static pressure fluctuation deviations in the nozzle and the test section at f = 6%. The effect of the nozzle shape at transonic flow velocities on the required main drive power is within the measurement error. CONCLUSIONS A new series of complex flow field investigations has been carried out in the TsAGI T-128 Wind tunnel (test section No 1). Main attention was paid to the influence of variable geometrical parameters (test section walls open area ratio, nozzle shape) on the time-averaged and non-stationary flow characteristics. Flow Mach and Reynolds number variations covered most of the operating range of the facility. Static pressure distributions were measured along the test section No 1 axis as well as on the walls and in the nozzle. Local flow inclinations were also measured on the test section axis. The results obtained show the improvement of the flow quality at the walls open-area ratio f = 6% compared with other f values. Local upwash is negative and does not exceed 0 1º in the absolute value. Velocity profiles, measured in the boundary-layer at the standard wall angle of inclination (θ = 0 5º), demonstrate the boundary layer thickness growth at small (f 1 5%) and large (f > 10%) open-area ratios. Pressure fluctuation levels were measured on the test section walls, in the nozzle and in the plenum chamber. The pulsation spectra are obtained and the main sources of perturbations are revealed. The compressor with numerous blades is the main source of the wide-band, low-frequency perturbations. In the highfrequency spectrum region, the tones caused by interaction of the free-stream flow with sharp perforation orifice edges are dominant. In terms of the number of tones generated and the amplitude uniformity of the acoustic spectra the optimum perforation ratio at transonic speeds is f = 6%. This value also gives the minimum rms static pressure fluctuation level in the test section and correlates with the local minimum of the facility-consumed power. Thus, 6% of the open-area not only ensures the minimum of wall interference and the best flow uniformity, but is also an optimum value from many other points of view. Change of the nozzle shape with the shift of the contraction zone 4-5 metres upstream creates a smoother flow transition from the nozzle to the test section and reduces the Mach number deviations

8 82 THE AERONAUTICAL JOURNAL FEBRUARY 2005 from the specified values by a factor of two. However, this is accompanied by noticeable boundary layer growth both at the nozzle exit and on the test section wall in the model location region. The consumed main drive power of the facility is not substantially sensitive to the nozzle shape changes. ACKNOWLEDGEMENTS The authors are grateful to Burov, V.V. and Khozyaenko, N.N., as well as to the whole personnel of the T-128 Wind tunnel for assistance in conducting tests and data processing. REFERENCES 1. NEYLAND, V.M., IVANOV, A.I., SEMENOV, A.V., SEMENOVA, O.K. and AMIRJANZ, G.A. Adaptive-wall perforated test section for transonic Wind-tunnels, AGARD-CP-585, 1997, 16, pp NEYLAND, V.M., IVANOV, A.I. and PILIUGIN, A.V. Some features of the test procedure in the new test section No 3 of the TsAGI Wind-tunnel T-128, Proc of the 2nd International Conference On Experimental Fluid Mechanics, 4-8 July 1994, Torino, Italy, pp GLAZKOV, S.A., GORBUSHIN, A.R., IVANOV, A.I. and SEMENOV, A.V. Recent experience in improving the accuracy of wall interference corrections in TsAGI T-128 Wind-tunnel. Progress in Aerospace Sciences, Pergamon Press, 2001, 37, pp HOLTHUSEN, H. and KOOI, J.W. Model and full scale investigations of the low frequency vibration phenomena of the DNW open jet. AGARD-CP-585, 1997, 26, pp STANNILAND, D.R.. MCHUGH, C.A. and GREEN, J.E. Improvement of the flow quality in the ARA transonic tunnel by means of a long cell honeycomb. Wind tunnels and Wind-tunnel test techniques. RAeS Conference, 1992, pp