experimental and numerical investigation of the cap-shock
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FR0301711
West-East High Speed Flow Field 2002 (WEHSFF) Conference Marseille (France), April 22-24, 2002
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Experimental and numerical investigation of the cap-shock structure in overexpanded thrust-optimized nozzles
Analyse experimental et numerique de la structure de choc en chapeau dans des tuyeres surdetendues a profit parabolique
par
P. Reijasse, F. Bouvier, P. Serve/
West-East High Speed Flow Field 2002 (WEHSFF) Conference Marseille (France), April 22-24, 2002
Resume : Ce papier traite de I'adrodynamique dans une tuyere surdetendue quend le profit parabolique interne du divergent de la tuyere est fortement optimise pour le rapport pouss6e/masse. Cette optimisation induit un choc interne qui se crde en aval du col de la tuyere meme si celle-ci fonctionne d des conditions proches du vide.Quand une telle tuyere est surdetendue le choc refiechi interfere avec le choc de surdetente et cette interference de chocs forme une structure particuliere appeiee structure de choc en chapeau en raison de la forme lumineuse ressemblant a un chapeau observee dans les sillages de tuyeres rdelles.Des calculs Navier-Stokes realises en Europe ont permis une analyse numerique d'un tel ecoulement et ont montre notamment une bulle de recirculation sur I'axe de la tuyere en aval du disque de Mach qui n'avait jamais pu etre mesurde jusqu'a present. Une campagne d'essais caractdrisant le ddcollement dans des tuyeres surdetendues a ete rdalisde dans la soufflerie R2Ch du centre de Chalais-Meudon a I'Onera. Des visualisations par strioscopie du jet en sortie de tuyere ont permis une description detaillde de la structure de choc en chapeau. Des mesures par veiocimdtrie laser a deux composantes ont confirmd I'existence d'une bulle de recirculation entourde d'un jet annulaire supersonique et ont fourni sa taille. En complement des calculs et des schemas interprdtatifs des strioscopies cette premiere caractdrisation quantitative experimental a conduit a une description physique du phenomena dans les tuyeres surdetendues.
NB : Ce document comporte 8 pages
Ce Tire a part fait reference au Document d'Accompagnement de Publication DAFE0208
West East High Speed Flow Fields 2002 D. E. Zeitoun, J. Periaux, J. A. D6sid£ri and M. Marini (Eds.)
© CIMNE, Barcelona, Spain 2002
EXPERIMENTAL AND NUMERICAL INVESTIGATION OF THE CAP-SHOCK STRUCTURE IN OVEREXPANDED THRUST-OPTIMIZED NOZZLES
Reijasse P.*, Bouvier F. and Serve! P.Office National d'Etudes et de Recherches Aerospatiales (ONERA), Fundamental and Experimental Aerodynamics Department,BP 72, 29 avenue de la Division Leclerc, 92322 Chdtillon Cedex, France e-mail: [email protected] web page: http://www.onera.fr
Abstract. This paper deals with the aerodynamics of an overexpanded-nozzle, when the internal parabolic contour of the nozzle extension is highly thrust-optimized in terms of specific impulse-to-weigth ratio. This optimization leads to an internal focusing shock issuing from a little downstream from the throat, even when the nozzle is running at nearly vacuum conditions. When such a nozzle is overexpanded, the focusing shock thus interfere with the overexpansion shock, and it forms from this shock interference a particular shock system, named ’’cap-shock” because of the cap-like luminous shape seen in the overexpanded plumes of some real engines. Navier-Stokes calculations performed in Europe had permitted to numerically analyze such a flow pattern, and they have revealed notably a recirculation bubble on the centerline downstream of the Mach disk, which had never been measured yet. A test campaign characterizing the flow separation in overexpanded subscale nozzles has been performed in the R2Ch blowdown wind tunnel of the Onera Chalais- Meudon center. Schlieren photographs of the exhaust jet have authorized a detailed description of the cap-shock pattern. Two-components Laser Doppler Velocimetry measurements have confirmed the existence of a recirculation bubble surrounded by an annular supersonic jet and has given its size. In addition to the calculations and the schlieren interpretative sketches, these first quantitative experimental characterization of the cap-shock structure permit to state a physical description of the cap-shock induced flow field in the thrust-optimized nozzles.
Key words: overexpanded nozzle, shock pattern, flow separation, Mach disk, wind tunnel.
1 INTRODUCTIONDuring the start-up of liquid rocket engine as SSME or Vulcain,1 * one can see dif
ferent visible structures in the jet downstream of the nozzle, according to the nozzle pressure ratio. One structure reveals the classical Mach disk as seen in Fig.l, but a
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K Keyasse et al./ Dap-shock structure m overexpanded nozzles
second one, which gives a cap-like luminous shape, also appears (see Fig.2).As nozzle side loads during engine start-up are mainly connected to the overexpanded flow properties in the nozzle jet, this paper aims to study this cap- shock structure, notably through Navier-Stokes computations2 and cold gaz subscale nozzle model experiments.3
Figure 1: Vulcain, overexpanded flow with Figure 2: Vulcain, overexpanded flow withclassical Mach disk (Courtesy of SNECMA) cap-shock pattern (Courtesy of SNECMA)
2 NOMENCLATURE
L divergent length from the throat to the nozzle exit 1, interaction lengthNPR nozzle pressure ratio (pst/pa)p stagnation pressure measured by Pitot tube p0 ambient pressure (in plenum chamber)Pe static pressure at the nozzle exitPs pressure at the separation locationPst stagnation pressurePo pressure at the origin of the shock/boundary layer interaction region Ue freestream velocityX abscissaY vertical axisZ longitudinal axis
3 EXPERIMENTAL SET-UP3.1 Wind tunnel
The tests have been performed in the Onera R2Ch wind tunnel (see Fig.3) of the Fundamental and Experimental Aerodynamics Department. This facility is a blowdown wind tunnel which can be used from Mach 3 to 7. The run duration
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f. Keyasse et al./ (Jap-shock structure in overexpanded nozzles
can reach several minutes because for these tests the tunnel is connected to the atmosphere. Otherwise test duration is about 30s when the tunnel is connected to the vacuum sphere.
Figure 3: R2Ch blowdown wind tunnel
3.2 Thrust-optimized contoured (TOC) nozzleThe axisymmetrical sub-scale nozzle model has an internal parabolic contour.
The main feature of this model is its ability to produce a thin annular supersonic layer at the wall which simulates the aerodynamic effect of a film cooling (see Fig.4). The nozzle consists of several parts numbered in Fig.4 :1. a converging part,2. a first diverging part of the nozzle with a 20mm-diameter throat,3. a second diverging part with a 112,9mm-diameter exit section,4. four secondary air injection chambers, each containing a throat of 1.8mm- diameter, a 5mm-diameter tube distributing the air mass flow through 4 rows of 6 holes of 2mm-diameter. The assembly of the two diverging parts of the TOC nozzle defines an annular injection slot with a critical section of 0.58mm-height. The injection slot is preceded by a settling chamber. In the diverging part of the nozzle, 3 generating lines (0 = 0°, 120° and 240°) are equipped with 19 continuous pressure taps. Figure 5 shows a photograph of the TOC nozzle model.
Figure 4: Sketch of the TOC nozzleFigure 5: TOC nozzle in the Onera R2Ch blowdown wind tunnel
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F. Keijasse et al./ Lap-shock structure in overexpanded nozzles
4 CALCULATIONSThe present calculations have been performed by using the NASCA research code.
This code is based on a Reynolds averaged Navier-Stokes (RANS) method.2 The figure 6 shows the results obtained with the Baldwin-Lomax turbulence model. The Mach isocontours are presented in the upper part of the figure as in the lower part, the streamlines are drawn in blue and the iso-pressures in red. This computation has been done for the TOC nozzle with an injection film and a nozzle pressure ratio Pst/Po equal to 50.
The most outstanding result is the recirculating bubble seen on the nozzle cen-
Mach: 0.0 0.5 1.0 1.5 2.0 2.5 3.0 3.5 4.0 4.5 5.0 5.5 6.0 6.5 7.0
Figure 6: RANS calculation of the Onera Thrust-Optimized Contoured nozzle (Nasca code) : iso-Mach field (top), isobaric lines (red bottom), streamlines (blue, bottom)2
terline of the Mach disk, which confirms other computational results performed in Europe2,4,5’.6 We clearly see that the subsonic region dowstream the Mach disk is characterized by the deviation of the streamlines which pass round the recirculating bubble. The huge subsonic region is surrounded by an annular supersonic flow submitted to successive compressions and expansions.
5 MEASUREMENTS 5.1 Schlieren visualisations
Figure 7 shows a schlieren picture of the jet issuing from the TOC nozzle at NPR=61. At this nozzle pressure ratio, the particular shock structure induced by the overexpansion regime in the TOC nozzle becomes visible. This structure has been called cap-shock pattern because of the cap-shape of the induced luminous regions observed in the real TOC nozzle jets.1 An interpretative sketch is given Fig.8.
This complex shock structure results from the interference between the internal -or focusing- shock (Cj) and the incipient overexpansion shock (Cj). The internal shock (C/) undergoes a reflection with a Mach disk, creating a triple point T% from which two other shocks emanate : the Mach disk (Cjv) and the reflected shock (Cm). A slip line (Si), very close to the reflected shock, is issuing from the triple point Ti. Such an angular divergence of the slip line from the centerline can be explained by
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f. Keijasse et al./ (Jap-shock structure in overexpanded nozzles
Figure 7: Schlieren photograph of the cap-shock Figure 8: Interpretative sketch of the cap-shock (Onera TOC-nozzle tested in R2Ch) structure
an inverse Mach reflection.7 A recent work8 has shown that the inverse Mach reflection could occur in permanent flow regime. We will see below that the deviation of the streamlines of the nozzle centerline permits the subsonic flow to slacken, then the static pressure downstream of the Mach disk to increase.The interference between the reflected shock (C#i) and the overexpansion shock (Cj) generates another Mach effect, with an annular shock (CU) and two triple points T2 and T3. Issuing from T3, the reflected shock (C«3) impinges the jet shear layer represented in non-viscid flow by the jet frontier or slip line (Sj). An expansion fan forms from this impingement, then reflects on another slip line (S3). Starting from T2, the reflected shock (C#2) impinges the slip line (Si), then successive expansion and compression waves, alternately reflect on (S2) and (Si). A small annular supersonic jet is bounded by these two slip lines, before vanishing rapidly. It results that this cap-shock pattern generates a huge supersonic flow region which forms downstream of the Mach disk (Civ). This subsonic zone is bounded by an annular supersonic jet, as shown in the schlieren picture of Fig. 7.
5.2 Pitot pressure probingsFor NPR close to 81, the normal shock is located 25mm downstream of the nozzle
exit plane, thus at X/L ~ 150mm. The measurement locations (X/L) and stagnation pressure ratios p/pit measured on the nozzle center line are given in the table 1.
Downstream of the shock, the pressure ratio p/pst is equal to 0.01227, which corresponds to the subsonic value MOIfs=0.395. Nevertheless the supersonic value Maxi, is also interesting because it corresponds to the value upstream of the normal shock. We notice that the static pressure downstream of the normal shock is a little less than the ambient pressure, paiw/Pa=0.9. The way for the pressure p downstream of the normal shock to rise up to the ambient pressure p0 is that the subsonic flow has to continue its deceleration. This can happen if the slip line (Si) issuing from the triple point is diverging from the nozzle centerline.
5.3 LDV measurementsThe LDV system has been used in its two-component version.9 The seeding of
the jet was ensured by silica oxyde particles injected far upstream of the model nozzle throat. The measurements were made in the freestream outer flow near the
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F. Keyasse et al./ (Jap-shock structure m overexpanded nozzles
Position of the Pitot probe relative to the Mach diskUpstream Downstream
X(mm) 122 155X/L (L=125mm) 0.976 1.240Pit/Po 80 82(p/Pst)axis 0.01870 0.01227
6.70 0.395Msup=7.38
Palis / Pa 1.331 0.896
Table 1: Pitot probings on the TOC nozzle axis
nozzle exit plane at various stations included in the area 25<X< 140mm from the nozzle exit and 0<Z<70mm. The streamlines and streamwise velocity contours are presented in Fig.9. Figure 10 gives a general view of the mean velocity vector field. The nearest station downstream of the nozzle exit is located at X=25mm. We can see that the streamwise component of the mean velocity becomes nearly equal to zero at X=75mm and can be even negative between X=75 and X=125mm in the area of the nozzle axis. A flow recirculation bubble has been detected in the region 75<X< 125mm and 0<Z<20mm. A probing along the nozzle axis, after the nozzle
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Figure 9: Streamlines and streamwise velocity Figure 10: Velocity profiles from LDV measure- iso-contours from LDV measurements ments
exit, confirms the size of the recirculation bubble, which is 50mm in streamwise direction (see Fig. 11).
5.4 Visualisations with a viscous coatingSome tests were performed with a viscous coating. Before the run, the lip of
the nozzle in the exit plane is covered by the coating. At the beginning of the run test, the coating is sucked in the nozzle and covered the nozzle walls. Then, with the shock issuing, a line appears and is recorded by a camera. Figures 12 and 13 provide the photographs for two nozzle pressure ratios (NPR=48.8 and NPR=63.1). The tests were performed for various nozzle pressure ratios from 47.2 to 85.9. For each of them, the location of the black line, which can be noticed in the previous figures, are reported with red lines in the graph of pressure distribution on the nozzle walls (Fig.14). A good correlation seems appear between these results. Nevertheless, according to Fig. 9, the front of the lines of the viscous coating seems not to represent
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P. Keyasse et ai./ Cap-shock structure in overexpanded nozzles
Figure 11: LDV probing in the nozzle axis from the exit plane
Figure 12: Photograph of the nozzle with viscous Figure 13: Photograph of the nozzle with viscous coating at NPR=48.8 coating at NPR=63.1
the separation line as expected. Maybe it can be explained by the viscosity of the coating which is not adapted. Further tests have to be performed.
itizzn wirnitieciwti cm
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Figure 14: Pressure distribution along the nozzle walls and "separation" lines location
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k*. Keijasse et al./ Cap-shock structure in overexpanded nozzles
6 CONCLUSIONThe cap-shock structure in overexpanded thrust-optimized contoured (TOC) noz
zles has been investigated in this paper, thanks to Navier-Stokes computations and to experiments performed in the R2Ch blowdown wind tunnel of the Onera Chalais- Meudon center.LDV measurements have confirmed the existence of a recirculation bubble surrounded by an annular supersonic jet, and have permitted, thanks to the help of schlieren photograph to state a physical description of the cap-shock structure in overexpanded nozzle jet.
ACKNOWLEDGEMENTS
The authors would like to thank ONES and SNECMA for their financial support. Special thanks are addressed to their colleagues of the "Flow Separation Control Device” program (Volvo Aero, Astrium, DLR, LEA at Poitiers) for their fruitful scientific exchanges.
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