( l, i AIAA 2002-0440 IMPROVEMENT OF FLOW QUALITY IN NAL CHOFU MACH 10 NOZZLE John Lacey Aero Systems Engineering, St. Paul, Minn, USA Dr. Yasutoshi Inoue NAL Chofu, Tokyo, Japan Akio Higashida Mitsubishi Heavy Industries, Kobe, Japan Manabu Inoue Mitsubishi Heavy Industries, Tokyo, Japan Dr. Kouichi Ishizaka Mitsubishi Heavy Industries, Takasago, Japan Dr. John J. Korte NASA Langley Research Center, Hampton, VA USA 40th AIAA Aerospace Sciences Meeting & Exhibit 14-17 January 2002 / Reno, NV For pennlmdon lo copy or republiah. ©erUct lha COl>Yrlghtowner named on litrefirm page, For AIAA-h4id c_y- dght, write to AIAA, Perml_dorm Department. 1801 AIm[ander Bell Drive, Suite 500, Reston, VA 20191-4344. AIAA - 2002-0440 IMPROVEMENT OF FLOW QUALITY IN NAL CHOFU MACH 10 NOZZLE John Lacey* Aero Systems Engineering, St. Paul, Minn, USA Dr. Yasutoshi Inoue** NAL Chofu, Tokyo, Japan Akio Higashida (Kobe), Manabu Inoue (Tokyo) and Dr. Kouichi Ishizaka (Takasago) Mitsubishi Heavy Industries, Japan Dr. John J. Korte** NASA Langley Research Center, Hampton, VA USA SUMMARY A program was initiated to improve the flow quality As a result of CFD analysis and remachining of the through the following steps: nozzle, the flow quality of the Mach 10Itypersonic 1) CFD modelling of the existing flow, Wind Turmel at NAL Chofu, Japan was improved. 2) Development of possible nozzle contours, The subsequent test results validated the CFD 3) Evaluation of these nozzle contours analytical predictions byNASA and Mt II. 4) Remachining of the nozzle, and 5) Validation testing. INTRODUCTION Mitsubishi Heavy Industries (MIII), Aero Systems The 1.27 meter diameter Mach 10nozzle atthe Engineering (ASE), and NASA Langley participated in this effort with NAL. MHI did the contour National Aerospace Laboratory (NAL) at Chofu, measurements and remachining; MHI and NASA LaRC Japan was designed inthe early 1990s for a nominal used their CFD capability to model the original test operating condition of 1070 K and 6MPa. This results and evaluate the new contours; and ASE nozzle was designed and manul_actured in accordance collaborated to develop the improved nozzle contours. with industry standards at the time. The flow quality (see Figure 1) of this project was acceptable at the ANAI,YSIS - EXISTING CONTOIJR time i;but since then, the standards of the industry have been tightened 2,and the Chofu design was not up to recent world-class perlbrmance criteria 3. The initial test data shown in Figure 1was expanded to include data at several other axial stations in the test section. As shown in Figure 2, the data indicate pressure waves traversing the test section and the effects of other waves converging oncenterline. The spread of Pitot pressure ratio isL 7%; this implies a Mach Number variation of+ 1.4% This profile was shown to be nearly symmetrical. Through a preliminary CFD analysis these waves were g)l 0 01 02 03 04 Radius Location -metens speculated to originate from the nozzle wall as shown in Figure 3 Figure 1Pilot Pressure Profile- Center of the Test Section ** Senior Member AIAA, * Member A[AA Copyright _c2,2002 The American Institute of Aeronautics and Asirollautics hlc, All rights rese_'ed 2 American Institute of Aeronautics and Astronautics (The differences that do exist are not deemed to be significant because aprediction of such steep gradients is subject to exact measurements of all peaks and I.i...... valleys and exact duplication of flow properties over the entire flow field. While the coordinates of the contour were measured in fine detail, there is the possibility that some peaks/valleys were not recorded and/or that some gas properties are not fully modeled.) 0003 00032 000.",4 0003G 00038 00o4 oo042 PitotPmtmlrl R.tlio¢X*fte_ Figure 2Pitot Pressure Profiles - Various axial 0.0033 stations _ 0.0032 . 2: ; .. _ _ V_ ' ...... D. .I 0.0031 0.0_0 -0.25 o o.25 0.5 y(m) ................, ......................................... 7*C,Oovn......................................... Figure 5 Comparisons of Pitot Pressure Profiles Figure 3Nozzle Sections and Possible Wave Pattern with CFD - NASA Langley While one wave could come from the joint between 16 Expansion 1and Expansion 2, the others were not so 9B ee obvious. Those others could be reflections of waves 94 that originated further upstream or they could be 9 waves generated on the nozzle wall at the location _) 02 04 ' 0'6 shown. Therefore, the wall contour was precisely =._ Radial Coordinales ,_._ _lal Co_¢_ates 06 measured at many stations so small disturbances • Meas_ernent -- CFDPredldKm lO 1o could be detected. The results of these measurements : -. • showed that the contour was, ingeneral, a shape typical of hypersonic nozzles but the slope and curvature were not fully smooth as shown in Figure 4 94 994 for the ttu'oatand initial expansion. 0'6 _ 02 o4 0'6 I_dial CoOrdlna_es Radl_ Coordinates Figure 6Comparisons of Mach Number Profiles with CFD - MHI OTHER TOOLS FOR ANALYSIS Along with the Pitot pressure profiles, the usual contour -100 100 300 500 700 900 1100 1300 1500 maps of Mach, pressure, velocity, etc. were also Axial Location -mm obtained (see Figure 7 for an example). The origins of Figure 4 Original slope and curvature waves were not easily interpreted from these usual contour plots and the parameter "GradP/P'" was The complete measured contour data was placed in developed. These contours are shown in for the original both the NASA-Langley _,__,a6nd MHI 7,8,9 CFD nozzle. As shown in Figure 8,many waves that come models and the predictions are shown in Figure 5and from the wall can be seen. Some ofthese waves get Figure 6respectively. In each case, the general cancelled through various interactions and a few waves character of the non-uniform profiles is predicted. enter the test section. Two of these waves generally 3 American Institute of Aeronautics and Astronautics agree with the two pressure waves that were identified low cost. The results are shown in Figure 9. Note that in in Figure 2. The strong gradients on the centerline are both the cases that were studied, major non-uniformities also obvious in this Figure 8. Contours could be still remain. In this case, it does not seem possible to analyzed more easily on the basis of this CFD achieve uni form flow with modification of only the parameter. throat block. IC(,:_4-:, ':-'-'(' 1--":" ._..,,,'<..'.,, "-, . 0 m_ 1 D. i T_N _wi t it fliltil l/)li I ol all o ll_i all 2 Nozzle 4Ler_ (m) 6 8 I xdl_l Oit llln ¢t-,_ Illlrli Figure 7 Typical Mach Contour Figure 9 Pitot Pressure Profile for Two Cases of Throat Block Smoothing - Compared with Original :l Final Smoothing Further attempts to obtain uniform flow consider the entire nozzle, not just the throat block. The original contour was based on the Method of Characteristics plus Boundary Layer correction (MOC+BL) and was then smoothed using cubic spline methods. For the new coordinates, arevised smoothing method was used based on polynomial equations of nozzle radius as a function of nozzle axial location. The nozzle was divided into axial sections for the curve fit procedure. In each section a 5thorder polynomial ethquation was applied. The polynomials were of the 5 order since this allowed the necessary inflections of the curvature. (For reference, acubic equation has a linear second Figure 8 Contour (if GradP/P for Original Nozzle derivative, i.e. curvature, and use of several cubic equations would have a discontinuous third derivative, not the smooth charateristic needed for a nozzle.) The CONTOUR SMOOTIIING lengths of the sections were adjusted to provide agood fit of the existing coordinates. Five sections were Initial Smoothing needed as shown in Figure 10. Based on the successful validation of the CFD At the match point between each section, the models, new contours were developed in an attempt to coefficients were adjusted until there was a match for obtain more uniform flow. One of the criteria for nozzle Radius, Slope, and Curvature. The section from these new contours was that most of the existing about 2500 mm to 5500 mm is covered by a single nozzle had to be re-used in order to minimize cost. equation. This is a subtle but important point; any wave The sections of the nozzle are shown in Figure 3. •originating in this range goes directly to the core of the test section, virtually without attenuation. This section The first attempt was to smooth just the throat block; must be avery smooth and this is best provided by a that is, the initial expansion of the nozzle. This is a single equation. small section and could be made new for relatively 4 American Institute of Aeronautics and Astronautics the need to limit the size of slope variations, especially Curve FitSections in the compression mode. Note that the sharp 700 compression comer near 700 mm may have contributed to the problem on centerline (after several reflections); ...... for reference see Figure 8that shows several strong waves originating in this area. Comparilon of New Conlour with Original Machined Contour 0 2000 4000 60190 8000 Axial Localion -mm Figure 10Sections of Nozzle fiw Curve Fit [ Procedure Further, since the MOC+BL method is step-by-step process that uses assumptions and averages to compute the downstream points, it ispossible that the ! best coordinates might be slightly different than the computed values. Therefore, exact matching of the Figure 11 Change in Contour MOC _BL coordinates was not a priority in this smoothing process. It is more important to have a continuous third derivative (smooth curvature) than to MACHINING RESULTS exactly match the MOC_BL computed points. The results of the machining to the new contour are One criteria for this nozzle was that the existing shown in Figure 12. The contour was held to within surface had to be mactfined at least 0.5 nun so that the nearly +,- 0.025 mm of the design values. The resulting new contour could be made with confidence. This compression comers near thejoints at 2500 and 5800 allowance was needed forthe tolerances of machine mm are less than 0.01 deg. There is a step, about 0.03 setup, concentricity, and a full removal of the existing mm high, near 2500 and a similar gap near 4100 ram. coating. This was accomplished by rotating the new As will be shown later, the flow quality isquite good contour about a point upstream of the throat. (The despite these deviations. rotation was only 0.0005 radians; additional smoothing was done in the throat area to maintain the throat size.) r 01 _--_ --e--Throat BlOck • .Expansion 1 _ Expansion 2 NEW CONTOUR - 1 [3 ExpaInsJion 3 _ Expansion 4 =1 [ The above method was used to generate several nozzle contours that might improve the flow uniformity. Each contour was evaluated using the CFD already validated to the test data as mentioned above. ol 0 1003 2_00 3000 4(X_ 50(_ 6000 1000 8(300 The final selected change in contour isshown in AxialLocation -mm Figure 11.This shows the slight modification that was applied tothe original nozzle. Note that, exclusive of Figure 12 Results of Machining to New Contour the general downward trend (resulting from the rotation mentioned above), the contour change isonly TEST RESULTS about +0.25 mm. The peaks inthe line at about 3400 and 5000 mm indicate compression comers (as Pitot Pressure compared to the new contour) and they nearly match Several improvements were made in the measurement the origin of waves suggested in Figure 3. Ateach of apparatus prior toconducting the final tests. The these two locations, the slope ischanged byabout 1.6 number of Pitot probes was increased and the span mm/1000mm (~0.1 deg total change); this highlights extended sothat measurements are at 10mm transverse 5 American Institute of Aeronautics and Astronautics spacinagndthemeasuremennotwsextenidntothe length of nearly constant Mach number inthe NAI, boundalrayyerT.hetesdt atawastakenatmoreaxial tunnel isapparent. stationws,ith50mmincremenintstheaxiadl irection coverin7g50me,inboththehorizontaanldvertical planeisncludintgestswithoff-axipslaceme(natsgain Nozzleexitcross section bothintheverticaalndhorizontoarlientationTsh).e (1.27rn Dia.) dataissummariziendFigure13;notethatthespread ofPitoptressuwreithinthecore+350mmaround (0.6m Dia.) centerlinheasbeerneducefrdom+7.5%to+1.5%. ThisimplietshattheMachNumbevrariatioinsnow beenreducefrdom L1.5% to ±0.3%. Core flowinuniform flow Summary of Test Data After Modification 000360 , Og , " ' 2" ". _.-o 9_" Figure 14 Flow inclination angle distribution at i ooo32o-I---,'_ ', ' " ' i,-- nozzle exit cross section et,. _11 ,._.,/*OriginaIx=0Nozzle _. !r ;' 5 K ooo3oo -t 102 ........ 000280 I .0,5 -0.4 -03 -02 -01 00 01 02 0.3 04 05 E _ AEDC, Z=64m m Radial Dimn¢o -metors •NAL, Z=-10_nm _U 9_ o NAL, Z=0mm m X NAL, Z=10mm Figure 13 Final Test Results 96 I mIIm IIeIIIIIII_IIsIIII__ llmIl Flow Anele I oCfeNllAt_LrofPt::_ShAecngtJDerM odelSupRpootzatti::ln The flow angularity was measured using arake with 94 .... _ ' ' ' conical probes. The measurements cover 0.72 min 600 400 200 0 -200 -400 width with 7conical probes spaced 0.12 m. These X,TunnelStatDn _AxJmlD:_ct_n,mm tests were made in 10axial locations, every 0.1 m Figure 15 Centerline Mach Number Variation - from nozzle exit (X 0) to 0.9 m downstream, with Compared with AEDC Tunnel C (from Reference the rake in both the vertical and horizontal position, 10). totaling 130 points. With the rake in ahorizontal position, the distribution of flow angle was obtained in 7vertical locations (Z 0,_0.12, _0.24, +0.3 m) at CONCLUSION 3longitudinal locations (X= 0, 0.25, and 0.5 m), An international team has studied the flow quality of totaling 147 points (some duplicating the longitudinal the NAL Mach 10nozzle. The participants used both positions). experimental data and computational methods to identify concerns and to modify the nozzle contour. The The flow inclination angles were from-0.02 degrees results with the new contour show that the flow quality to 0.1 degrees on the centerline and within i0.2 has been significantly improved. As aresult the NAL degrees in the whole core flow region. An example of Mach I0 wind tunnel can now serve the aerodynamic the result at the nozzle exit cross section is shown in community as a world-class testing facility. Figure 14.The flow is nominally axisymmetric with slight inward flow at this location. ACKNOWLEDGEMENT Centerline Mach Number The authors wish toacknowledge the contributions of Further evidence of the flow quality achieved inthis the following people that also contributed to the project effort isshown in Figure 15(taken from Reference and this paper; namely Mr. Seizo Sakakibara and Dr. 10). This shows the comparison ofthe centerline Koichi Hozumi ofNAL Chofu. Mach number inthis wind tunnel as compared tothe same parameter inthe AEDC Tunnel C. The long 6 American Institute of Aeronautics and Astronautics REFERENCES 1) Boudreau, A.H., "Performance and operational characteristics of AEDC/VKF tunnels A, B, and C," AEDC-TR-80-48 (198l). 2) Micol, J. R., "ttypersonic Aerodynamic/Aerothermodynamic Testing Capabilities at Langley Research Center: Aerothermodynamic Facilities Complex," AIAA-95-2107. 3) Aerodynamics Department, "Plan and Structure of Large-size Hypersonic Wind "Funnel," Technical Report of NAL, TR-1261 (1995). (in Japanese) 4) Korte, J. J., Kumar, A., Singh, D. J. and White, J. A., "CAN-DO - CFD-Based Aerodynamic Nozzle Design & Optimization Program for Supersonic/tlypersonic Wind Tunnels," AIAA Paper 92-4009, July 1992. 5) Korle, J. J., Hedlund, E. and Anandakrishnan, S., "'A Comparison of Experimental Data With CFD For The NSWC Hypervelocity Wind Tunnel //9 Mach 14 Nozzle," AIAA Paper 92- 4010, July 1992. 6) Korte, J. J. and Hodge, J. S., "Flow Quality of tlypersonic Wind-Tunnel Nozzles Designed Using Computational Fluid Dynamics," J. of SpacecraB and Rockets, Vol. 32, No. 4, July- August 1995, pp. 569-580 7) lshizaka, K., Ikohagi, T. and Daiguji, H., "A High Resolution Finite-Difference Scheme for Compressible Gas-Liquid Two Phase Flows," Proc. of the 5th ISCFD, Vo1.1(1993), pp.352- 357 8) Eggers, Jr. A. J., "One-Dimensional Flows of an Imperfect Diatomic Gas," NACA Report 959, 1950 9) AMES Research Staff, "Equations, Tables And Charts For Compressible Flow," NACA TR 1135, 1953 10) Nagai, S., Tsuda, S., Koyama, T., ttirabayashi, N., and Sekine, It., "Comparison of Winged Vehicle Force Data at Large Hypersonic Wind Tunnels", AIAA-2001-0166. 7 American Institute of Aeronautics and Astronautics