Pressure coefficient evaluation on the surface of the SONDA III model tested in the TTP Pilot Transonic Wind Tunnel

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1 Jornal of Physics: Conference Series OPEN ACCESS Pressre coefficient evalation on the srface of the SONDA III model tested in the TTP Pilot Transonic Wind Tnnel To cite this article: M L C C Reis et al 015 J. Phys.: Conf. Ser Related content - Simlation of erosion by a articlate airflow throgh a ventilator A Ghenaiet - Planck-LFI radiometers tning F Cttaia, A Mennella, L Stringhetti et al. - Effect of Airflows on Reetitive Nanosecond Volme Discharges Tang Jingfeng, Wei Liqi, Ho Yxin et al. View the article online for dates and enhancements. This content was downloaded from IP address on 17/09/018 at 19:34

2 Jornal of Physics: Conference Series 588 (015) doi: / /588/1/01003 Pressre coefficient evalation on the srface of the SONDA III model tested in the TTP Pilot Transonic Wind Tnnel M L C C Reis 1, J B P Falcao Filho 1, E Basso 1 and V R Caldas 1 Institte of Aeronatics and Sace Pr. Mal. Edardo Gomes, 50, CEP 18904, São José dos Camos, SP, Brazil University of Tabate Av. 9 de Jlho, 199, CEP , Tabate, SP, Brasil marialisamlccr@iae.cta.br Abstract. A test camaign of the Brazilian sonding rocket Sonda III was carried ot at the Pilot Transonic Wind Tnnel, TTP. The aim of the camaign was to investigate aerodynamic henomena taking lace at the connection region of the first and second stages. Shock and exansion waves are exected at this location casing high gradients in airflow roerties arond the vehicle. Pressre tas located on the srface of a Sonda III half model measre local static ressres. Other measred arameters were freestream static and total ressres of the airflow. Estimated arameters were ressre coefficients and Mach nmbers. Uncertainties associated with the estimated arameters were calclated by emloying the Law of Proagation of Uncertainty and the Monte Carlo method. It was fond that both ncertainty evalation methods reslted in similar vales. A Comtational Flid Dynamics simlation code was elaborated to hel nderstand the changes in the flow field roerties cased by the distrbances. 1. Introdction Scale models of aerosace vehicles are tested in wind tnnels in order to redict the erformance of their fll scale conterarts dring actal flight. This stdy resents the reslts of a test camaign of a half-model of the SONDA III sonding vehicle carried ot at the Brazilian Pilot Transonic Facility, TTP. The rose of the tests is to evalate changes in the airflow field roerties de to the resence of shock and exansion waves in the inter-stage region of the vehicle. The evalated arameter is the ressre coefficient (-C ). Measred arameters are the total airflow temeratre, T 0, total and static ressres, 0 and, and static ressre,, taken at ressre tas located at stations distribted longitdinally along the model srface. Uncertainties in the measred arameters are evalated and roagated to the ressre coefficient by sing both the Law of Proagation of Uncertainty and the Monte Carlo method. Mach nmber of the airflow, M, and associated ncertainty are also evalated. Exerimental data are comared to nmerical data obtained throghot Comtational Flid Dynamics, CFD. 1 To whom any corresondence shold be addressed. Content from this work may be sed nder the terms of the Creative Commons Attribtion 3.0 licence. Any frther distribtion of this work mst maintain attribtion to the athor(s) and the title of the work, jornal citation and DOI. Pblished nder licence by Ltd 1

3 Jornal of Physics: Conference Series 588 (015) doi: / /588/1/ The sonding rocket SONDA III The aerosace vehicle SONDA III (figre 1) is a bi-stage vehicle develoed by the Institte of Aeronatics and Sace, IAE. It is one of the sonding rocket family named Sonda, which started with Sonda I, first lanched in In 1996, the first stage of Sonda III was evolved to receive Eroean exeriments onboard. This new single stage vehicle was known as VS-30. A boosted version, the VSB-30, which contains the S31 booster motor, was develoed in 001. This bi-stage geometric concetion has been tilized in many aerosace designs and a better nderstanding of the aerodynamic henomena taking lace arond the vehicle is imortant for ftre design initiatives. Figre 1: The Sonding Rocket Sonda III. With this rose, a half-model of the vehicle was tested in the TTP wind tnnel in order to stdy the flow field in the inter-stage region, where shock and exansion waves are exected. Shock waves arise on the vehicle srface when the airflow becomes locally sersonic. The flow roerties change throgh the shock wave and ndesirable effects which comromise the flight erformance may occr. Therefore, an investigation of the behavior of shock and exansion waves in wind tnnels is imortant to characterize their interactions with the model. The nominal Mach nmber vales covered in the tests varied from M = 0.0 to Shock and exansion waves When a rocket vehicle is in a free flight condition at high velocity, the flow may become locally sersonic over the srface of the vehicle [1]. The freestream Mach nmber at which M = 1 is first achieved on the srface is called the critical Mach nmber. In sersonic regimes, obliqe shock waves occr when sersonic flow enconters a concave corner and it is attached to the corner as long as the concave corner angle is below a defined limit. Above this limit a more comlex detached shock wave occrs. In contrast, if the sersonic flow enconters a convex corner, exansion waves occr. Flow field roerties change across shock and exansion waves, as shown in figres a and b. Tables and diagrams relating the angle of the corner (θ ), the angle of the shock (β), and the Mach nmber (M) can be fond in [].

4 Jornal of Physics: Conference Series 588 (015) doi: / /588/1/01003 Region 1: stream of the shock M 1 > 1 streamlines Region 1: stream of the exansion fan M 1 > 1 streamlines β concave corner θ a) Exansion fan Obliqe shock Region : downstream of the shock M < M 1 > 1 T > T 1 ρ > ρ 1 Region : downstream of the exansion fan M > M 1 < 1 T < T 1 ρ < ρ 1 θ convex corner b) Figre : Deflection of a sersonic flow in the resence of: (a) concave corner, (b) convex corner. M: Mach nmber., T, and ρ: static ressre, temeratre and density, resectively The TTP wind tnnel The Pilot Transonic Wind Tnnel, TTP, is located at the Aerodynamics Division of the Institte of Aeronatics and Sace, Brazil (figre 3). The tnnel is continosly driven by an 830 kw main axial comressor and can also be oerated intermittently by means of an injection system, which slies airflow for arond 30 seconds. It is a variable-ressre wind tnnel with control caability to indeendently vary Mach nmber, stagnation ressre, stagnation temeratre and hmidity. The test section is 0.5 m high and 0.30 m wide and has longitdinally slotted walls to favor the niformity of the airflow. At TTP, the Mach nmber range can vary from 0.0 to 1.3. Some configrations can be changed in order to allow better flow control sch as reentry flas osition and the rate of forced mass extraction, determined by the Plenm Evacation System, PES. Figre 3: The TTP aerodynamic facility. 3

5 Jornal of Physics: Conference Series 588 (015) doi: / /588/1/ Descrition of the Model Figre 4 shows a 1:8 scale model concetion, monted on an Alminm latform which will be fixed to the lateral wall of the tnnel test section. In figre 4 the model is configred with one fin erendiclar to the tnnel wall. Threads made in secial laces allow the model to change the direction in relation to the latform in order to adjst angles of attack of 0 o, 5.5 o, 8 o, 1 o, 13 o and 15.5 o. Between the half-model and the latform secial sacers can be installed to create a ga between the tnnel wall and the model. This way the ossible bondary layer formed on the tnnel wall can be eliminated and ths better reresents the free flight sitation. The total length of the model is m and the diameter of the first stage is m. Figre 4: Model concetion of Sonda III with scale 1:8 monted on an Alminm latform. The Sonda III model with details of lastic tbes is resented in figre 5. The tbing connects the ressre tas and ressre instrmentation. There are a total of 154 measring ressre tas located in ositions along the model srface. The ositions are identified by letters A to V and are distribted in a region encomassing the inter-stage frstm cone (figre 6). The classification is in alhabetic order, dearting from the base of the first stage in the direction of the nose of the second stage. Tas A to H belong to the first stage, this latter defining the limit border of the connection between both stages. The connection region has a frstm cone shae, with tas I, J, K and L. The remaining ressre tas, M to V are located in the second stage of the model. Figre 5: Sonda III model with detail of the tbe connections. 4

6 Jornal of Physics: Conference Series 588 (015) doi: / /588/1/01003 Figre 6: Pressre ta ositions arond the inter-stage region.. Metodology This stdy describes reslts obtained with the half-model installed on the test section lateral wall of the tnnel. The aim of the wind tnnel tests is to determine the ressre distribtion on the model srface. This section describes the data redction of the estimated arameters. Firstly, the mathematical models of the estimated qantities, which are the negative ressre coefficient, -C, and the freestream Mach nmber of the airflow, M, are resented. Afterwards, the Law of Proagation of Uncertainty is alied to the mathematical models of both arameters, to estimate the associated ncertainties. Information abot the Monte Carlo imlementation is then given, sch as the nmber of trials and robability distribtions of the measred qantities. The section ends with a descrition of some roerties of the Comtational Flid Dynamics code emloyed in the nmerical simlation..1. Pressre coefficient The ressre distribtion on the model srface can be reresented in terms of the ressre coefficient defined by []: q C P (1) where is the local ressre measred on each ressre ta station of the model, is the freestream static ressre and q is the freestream dynamic ressre. The static ressre sensor sed to measre is ositioned on the er art of the wall at the beginning of the wind tnnel test section. The freestream dynamic ressre, q, is defined by: q 1 ρv () where ρ and V are the freestream density and velocity, resectively. In this stdy, the air is considered as a erfect gas. Its density is calclated by: = RT ρ (3) : freestream ressre in ascal; R: erfect gas constant, eqal to 87 J/(kg.K) for normal air; and T : freestream temeratre exressed in kelvin. The airflow Mach nmber, M, is the ratio between velocity, V, and the seed of sond, a: M = V/a (4) 5

7 Jornal of Physics: Conference Series 588 (015) doi: / /588/1/01003 In eqation (4), a reresents the seed of sond travelling throgh the air, considered in this stdy as a erfect gas. Its vale is estimated by: a = γrt (5) where T: temeratre exressed in kelvin; and γ: ratio of secific heats, eqal to 1.4 for air considered as a erfect gas. Rearranging eqation (1) by sing (), (3), (4) and (5), reslts in: C P = 1 γ M (6) In regions of the flow where >, the ressre coefficient -C will be a negative vale. Considering the flow in the test section as comressible flow, the freestream Mach nmber, i. e., the Mach nmber at ndistrbed condition (before reaching the model) is given by: M = γ 1 ( γ 1) γ 0 1 (7) where 0 is the freestream total ressre. As already mentioned, a static ressre ta is located on the er wall of the test section and connected to a ressre sensor to measre. The total ressre sensor sed to measre 0 is located in the stilling chamber of the circit... Law of roagation of ncertainty According to reference [3], the vale of the ncertainty in measrement is the ositive sqare root of eqation (8): c = N i= 1 y x i ( x ) i (8) where y is the ott qantity and x i s are the int qantity. Eqation (8) is called the Law of Proagation of Uncertainty and its evalation is known as the GUM ncertainty framework [4]. Alying eqation (8) to (6) reslts in the vales of the ncertainty in the ressre coefficient: C = C C + C + M M (9) which leads to: 6

8 Jornal of Physics: Conference Series 588 (015) doi: / /588/1/01003 C = + + M γ 3 M γ M γ M 1 ( ) (10) In eqation (10), the ncertainties in the static ressre,, and in the freestream ressre,, are estimated combining two comonents: the ncertainty declared in the calibration certificate of ressre instrmentation and the ncertainty evalated as tye A, qantified by calclating the standard deviation of the temoral ressre signal. To evalate the ncertainty in the freestream Mach nmber, one alies the Law of Proagation of Uncertainty to eqation (7): M M = 0 M + 0 (11) which reslts in: = + M 0 M 0 0 (1).3. Monte Carlo method Reference [4] was sed to sly Mach nmber and ressre coefficients vales and associated ncertainties by sing the Monte Carlo method. The int qantities for the freestream Mach nmber, M, are the total ressre, 0, and the freestream static ressre, (eqation 7). The int qantities for the ressre coefficient, -C, at each ressre station on the half SONDA III model, inclde M itself and the local static ressre taken at ressre tas A to V, along with 0 and of the airflow (eqation 6). The mean vale and standard deviation of the ressre signals slied by the instrments dring the rns of the wind tnnel tests are comted and this information is sed to evalate the measrement ncertainty by sing the Monte Carlo method. The robability distribtions for the int qantities are roagated throgh the measrement models exressed by eqations (6) and (7), reslting in the estimated ott qantities -C and M and the associated standard ncertainties, -C and M (eqations 10 and 1). Codes in MatLab were develoed for the imlementation of the Monte Carlo method. Gassian distribtions were assigned for all int qantities. The selected nmber of trials was 50, Comtational Flid Dynamics code The nmerical code sed to solve Reynolds-Average Navier-Stokes eqations was the CFD++ solver. The wind tnnel walls were not considered in the simlation therefore the vehicle is in free flight condition. In this stdy, the model reresentation is confined to a dihedral angle of 10 o, with origin at the symmetric axis of the model fselage. This is ossible becase the simlated model is withot fins and is ositioned at nll angle of attack. These conditions lead to axisymmetric flow and it is ossible to limit the comtational field. Pre-rocessing was carried ot by sing ANSYS ICEM CFD meshing software. The airflow is considered viscos and the trblence model is the Salart Allmaras [5]. The reslting y+ vale is less than 1, adeqate for the investigation of bondary layer effects. Meshing is comosed of hexahedrons and risms. The code alication for this kind of rocket geometry was validated by sing a cone-cylinder model. 7

9 Jornal of Physics: Conference Series 588 (015) doi: / /588/1/ Reslts and discssion The test camaign to analyse the ressre coefficient distribtion on the Sonda III half-model srface carried ot at the transonic wind tnnel TTP considered the nominal Mach nmbers 0.0, 0.30, 0.40, 0.50, 0.60, 0.70, 0.80, 0.85, 0.90, 0.95, 1.00, 1.05 and The angle of attack of the model is 0 o. The ressre signals obtained for total ressre, 0, and freestream static ressre,, as a fnction of time, are shown in figres 7a and 7b resectively. The resented signals were obtained for nominal Mach nmber, M, eqal to When the desired regime settles down, an interval of the ressre signals is chosen and the average and standard deviation are calclated. The interval between 40 s and 10 s was chosen for M = 0.95, and the transient of the signal discarded. The transient is cased by the wind tnnel drier system, which controls the hmidity of the airflow. The ratio 0 / of average vales of total and static ressres is comted. This ratio is inclded in eqation (7) to sly the freestream Mach nmber vale, M. The same rocedre is erformed for static ressre vales in tas A to V in order to obtain ressre coefficients (eqation 6). a) b) Figre 7: Temoral ressre signals for: (a) total ressre, (b) freestream static ressre. In Figre 8 one can see the local Mach nmber at each ressre ta of the model, for all regimes covered by the wind tnnel tests. The origin of the x-axis coincides with the beginning of the frstm cone and is located between tas L and M. For nominal Mach nmber 0.80 there is a region on the model srface where the local Mach nmber is above 1, indicating that the critical Mach nmber is between regimes 0.70 and Figre 8: Mach nmber at ressre tas A to V. 8

10 Jornal of Physics: Conference Series 588 (015) doi: / /588/1/01003 Figre 9 and table 1 resent the ressre coefficient vales for nominal Mach nmber regime 0.90, obtained at ressre tas ositions A to V of the SONDA III model, by sing eqation (6). The airflow direction is from left to right in the ictre. Uncertainties estimated by sing the GUM aroach (eqation 10) are comared to the Monte Carlo reslts in table 1. No differences were fond when comaring both methods. Pressre coefficient data are resented in table for regimes 0.60 and Only Monte Carlo reslts are shown WIND exerimental nc nc -C P ressre ta osition (mm) Figre 9: Pressre coefficient and associated ncertainty at ressre ta locations. M = Table 1: Pressre coefficients and associated ncertainties (vales were mltilied by10). M = Pressre ta A B C D E F G H I J K -C GUM GUM C -C MC MC C Pressre ta L M N O P Q R S T U V -C GUM GUM C -C MC MC C 9

11 Jornal of Physics: Conference Series 588 (015) doi: / /588/1/01003 Table : Comarison of ressre coefficients and associated ncertainties for sbsonic and sersonic regimes. Pressre ta A B C D E F G H I J K M = C C Pressre ta L M N O P Q R S T U V -C C M = 1.10 Pressre ta A B C D E F G H I J K -C C Pressre ta L M N O P Q R S T U V -C C A qantitative comarison between Comtation Flid Dynamics reslts and exerimental data is lotted in figre 10 for nominal Mach nmber The nmerical and exerimental ressre distribtions agree to a certain extent, bt considerable differences occr at oints E and F. The deartre from exerimental data reveals that the simlation was not adeqately caable of catring details of aerodynamic henomena occrring in this region. Uncertainty limits are not sfficient to exlain the discreancies between exerimental and simlation data. Differences may have been cased de to the fact that the simlation considers the entire model, not the half-model attached to the wind tnnel wall. Therefore, the simlation is for a vehicle traveling in free flight. Half-model comtational simlation wold be desirable exerimental nc nc simlation C ressre ta osition (mm) Figre 10: CFD and exerimental reslts for freestream Mach nmber eqal to

12 Jornal of Physics: Conference Series 588 (015) doi: / /588/1/01003 Pressre coefficients obtained for all Mach nmber regimes are lotted together in figre 11. Uncertainty vales were omitted for clarity. Vertical axis reresents -C. A schematic reresentation of the model region where the measrements were taken was also drawn at the bottom of the grah, highlighting the initial and the final sections of the frstm cone. Following the wind direction in figre 11, in the region of ressre tas of the second stage identified by T, U and V, the static ressre is very close to, reslting in near zero vales of -C (eqation 6). This fact reveals that aerodynamic distrbances were not relevant at these oints. Nevertheless, one can note a different behavior between sbsonic and sersonic airflows, i. e., sersonic Mach nmber ressre distribtions are aart from the sbsonic ones. This also occrs after the frstm cone, in ositions A, B and C. Enlarged versions of these regions are shown in figre 1. Figre 11: Pressre coefficient distribtion on model srface. M = 0.0 to a) b) Figre 1: Regions resenting ressre coefficient vales near zero: (a) second stage, (b) first stage. To better visalize the effects of the concave and convex corners in the inter-stage region and the resence of shock and exansion waves, grahics are ordered in increasing vales of Mach nmber in figre

13 Jornal of Physics: Conference Series 588 (015) doi: / /588/1/01003 a) b) c) d) e) f) Figre 13: Comarison of -C distribtions for increasing Mach nmber. One can see similar ressre distribtion atterns for the sbsonic ranges in figre 13a for Mach nmber to As a reslt of the distrbance cased by the concave corner arond oint L, the airflow static ressre starts to increase and reaches its maximm vale at this oint. From oint L to I, which are located in the frstm region, the airflow local static ressre continosly decreases. This tendance remains to ressre ta H, giving the minimm ressre vale. From this oint on, the ressre starts to recerate the freestream condition. The attern is modified from figre 13b on. The local ressres after the area exansion change in sch a way that the ressre coefficient distribtion becomes more comlex. From figre 13c on, the regime is sercritical. In figres 13c and 13d, it is ossible to observe the sccessive growth of -C de to the exansion region arond the convex corner at oint H. The difference ( -) becomes more ositive for Mach nmber eqal to 0.90 and -C achieves maximm vale. The -C did not contine to increase, as shown in figre 13e. At freestream sersonic flow condition, a shock wave wold be exected in the region of the ressre ta L de to the concave corner if the relation θ β M (angle of the corner, angle of the shock and Mach nmber) was adeqate. The local static ressre behind the shock increases in relation to, reslting in negative vales for -C. For Mach nmbers 1.05 and 1.10, the airflow resents this 1

14 IMEKO Jornal of Physics: Conference Series 588 (015) doi: / /588/1/01003 behavior even before reaching the concave region (see figres 1a and 13f). Aarently, the shock wave is detached de to the high angle of the frstm cone, and negative vales start to occr earlier, characterizing a local sbsonic regime. Simlation by emloying comtational flid dynamics was done for the Mach nmber regime Some roerties of the nmerical code have already been mentioned in section.4. A diagram of the whole comtational field defined by a dihedral angle of 10o is shown in figre 14, highlighting the central lan. The model is ositioned at the botton right and the distance is defined from the begining of the first stage of the model. Figre 14: Diagram of the airflow field; x-axis nit: meter. The comtational grid toology arond the vehicle for the central lane is shown in figre 15a. Note the grid clstering where a better anlysis of areodynamic henomena is desired, i. e., inter-stage region, nose and next to the srface. A detailed view of the oint clstering at the inter-stage art can be seen in figre 15b. The very closely saced grid in the vicinity of the model srface is necessary in order to accratelly catre the bondary layer effects, according to the reqirement (y+ < 1) of the Salart Allmaras trblence model [5]. There are 10 circmferentially and niformly saced lanes (1o between lanes). The whole field has 48,760 calclation cells. a) b) Figre 15: a) Comtational meshing arond the vehicle. b) A zoom view of the inter-stage region. 13

15 Jornal of Physics: Conference Series 588 (015) doi: / /588/1/01003 Figre 16 smmarizes the henomena which occr arond the model at an angle of attack eqal to 0 o in Mach nmber regime The airflow arameter shown is static ressre. The freestream ressre is eqal to 55,580 Pa. The flow is from left to right. One can observe the static ressre increasing before the frstm cone and the ressre decreasing after this region. After assing throgh the cone, exansion waves are resent followed by a shock wave which is identified by the coalescing of isobaric crves. This shock wave interacts with the bondary layer in the vicinity of the model srface reslting in shock thickening. The bondary layer grows after the assage of the shock wave, becase of the adverse ressre gradient. This fact is better seen in the Mach nmber field shown in figre 17. The sersonic region defined by the line corresonding to Mach nmber 1.00 starts almost at the end of the frstm cone and has an extension of 0.37 times the diameter of the first stage. Figre 16: CFD static ressre in the inter-stage region. M = 0.90 Figre 17: Mach nmber field. M = Figre 18 is a detailed view of the shock-wave bondary-layer interaction region with velocity vectors. One can see the redction of velocity rofile (region in ble), cased by the shock wave iminging over the bondary layer. As a conseqence, the bondary-layer external frontier is enlarged. 14

16 IMEKO Jornal of Physics: Conference Series 588 (015) doi: / /588/1/01003 Figre 18: A detail of the velocity rofile. 4. Conclsions The aer has resented the reslts for a test camaign carried ot at the Pilot Transonic Wind Tnnel. A half-model of the sonding rocket SONDA III, eqied with ressre tas arond the inter-stage region, was tested in the sbsonic and transonic regimes. Pressre coefficient distribtions arond the vehicle were estimated, as well as the associated ncertainties, to sly information abot aerodynamic henomena taking lace in the airflow arond the vehicle. Uncertainties were evalated by sing the Law of Proagation of Uncertainty and Monte Carlo method. It was fond that both methods reslted in similar ncertainty vales and these were not significant to indicate the resence of measrement distrbances cased by the aerodynamic circit. A Comtational Flid Dynamics nmerical code was sed to better nderstand the changes in airflow roerties cased by the resence of the model. This aer detailed the simlation reslts for Mach nmber eqal to Exerimental and simlation data did not agree comletely and cases will be better investigated in ftre stdies. 5. References [1] [] [3] [4] [5] Délery J and Dssage J P 009 Some hysical asects of shock wave/bondary layer interactions Shock waves (Sringer-Verlag) DOI /s z Anderson Jr J D 007 Fndamentals of Aerodynamics (Mc Graw Hill) New York 985 BIPM/JCGM 100:008 Evalation of measrement data Gide to the exression of ncertainty in measrement (GUM 1995 with minor corrections) 134 BIPM/JCGM 101:008 Evalation of measrement data Slement 1 to the Gide to the exression of ncertainty in measrement Proagation of distribtions sing a Monte Carlo method 8 Wilcox D C 1993 Trblence Modeling for CFD La Cañada DCW Indstries