An Experimental And Numerical Investigation Of Supersonic Contour Nozzle Flow Separation

Proceedings of the International Conference on Aerospace Science and Technology 26 - 28 June 2008, Bangaloire, India

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INCAST 2008- 074

AN EXPERIMENTAL AND NUMERICAL INVESTIGATION OF SUPERSONIC CONTOUR NOZZLE FLOW SEPARATION

1

M. Venkata Raman1, C. Senthil Kumar2 and Dr. S. Elangovan3

PG Student, Dept of Aerospace Engineering, MIT, Anna University, Chennai, India, [email protected]. 2 Lecturer, Dept of Aerospace Engineering, MIT, Anna University, Chennai, India, [email protected]. 3 Assistant Professor, Dept of Aerospace Engineering, MIT, Anna University, Chennai, India, [email protected].

ABSTRACT The project presents, preliminary experimental and numerical results on separation of supersonic flow inside a convergent-divergent (CD) contour nozzle. The study is motivated by the flow separation occurring inside CD nozzles operated at low pressure ratio. A novel apparatus allows investigation of many pressure ratios with large optical access and measurement of wall pressures. The separated flow in a convergent-divergent (CD) contour nozzle is investigated by the solution of the Reynolds-Averaged Navier-Stokes equations with a two-equation k-ε turbulent model. For a fixed area ratio, defined as the exit area to the throat area, Ae / At of 1.638, computations are conducted over a range of nozzle pressure ratio (NPR). The computational results are used to further examine the effect of the separated flow. Unlike the experiment, no unsteady shock movements are simulated. The computed wall pressure distribution is in good agreement with the experimental data. Also consistent with experimental observations, the flow separates asymmetrically for the range of 1.5
INTRODUCTION Flow separation may occur in an convergent-divergent (CD) contour nozzle due to the nozzle expansion ratio being too large for a given nozzle pressure ratio (NPR). In that case, supersonic contour nozzle flow expands to a pressure level that is far lower than the ambient pressure, causing the formation of a normal shock and detachment of the ensuing flow from the nozzle walls. For some nozzle shapes, the flow separation can generate large unpredictable sideloads. The survey shows that the separation as an undesirable occurrence, and showed that off-design nozzle thrust efficiency could be greatly improved by controlling the location and extent of the separation. The understanding and prediction of the physical mechanisms behind the supersonic nozzle separation flow is important for designing efficient nozzle configuration that avoids asymmetric side forces and enhances flow mixing. In that case, supersonic (CD) contour nozzle design is carried out using Method of characteristics, region method. Although there is a large amount of literature published for separated nozzle, the detailed investigation on separation flow mechanism is less matured. Most of available publications are concentrated on prediction of separation location. A review paper by Morrisette and Goldberg, based on a variety of experimental results concludes that zero-pressure gradient separation prediction method, like that proposed by Reshotoko and Tucker, give reasonable predictions for nozzles with turbulent separation and large divergence angle.

Proceedings of the International Conference on Aerospace Science and Technology 26 - 28 June 2008, Bangaloire, India

International Conference on Aerospace Science and Technology 26-28 June2008, Bangalore, India

In the paper of Romine, the mechanism for causing the flow separation from the nozzle wall is demonstrated, and the theory for a new solution of the separation location is presented. Near the center-line of the nozzle, Romine postulates that the pressure adjusts to the ambient pressure via a gradual under expansion. On the wall, the flow adjusts to the ambient pressure almost immediately past the shock. Some previous studies have conducted numerical prediction of separated nozzle flows and the internal performance, such as Hunter, Carlson, Xiao. In general, the computation is in excellent agreement with experimental data. In the paper of Hunter, two distinct separation regimes are shown from their experimental results on a two dimensional planar nozzle with Ae/At =1.8. For NPR<1.8. The flow shows threedimensional separation with partial reattachment. Fully-detached two-dimensional separation is found for NPR>2.0. The under expansion of flow after the main shock is evident from the computed Mach number contours, although this was not explicitly expressed. Recently, Papamoschou and Zill11 investigated experimentally the supersonic separated flow inside a symmetric convergent-divergent nozzle.

DESCRIPTION The asymmetric flow separation occurring in geometrically symmetric nozzle has not been computed. For experimenting the contour CD nozzle is fabricated using Method of Characteristics. The nozzle made of Hardened Stainless Steel using Electric Discharge Machining. This has 16 pressure ports, which is made available to connect a digital pressure sensor (Scannivalve) for the accurate pressure measurement. The wall pressure is measured by 2000 frames per minute using the DSA Link software. Unsteady Reynolds-averaged Navier-Stokes equations with RNG, K-ε and K-ω turbulent model were used for the computation.

Fig. 1 shows the nozzle designed by Method of Characteristics and fabricated using Electric Discharge Machining.

RESULT & DISCUSSION An investigation has been conducted into the source of plume instability from convergent-divergent contour nozzles. The effect of internal shock phenomena on the plume unsteadiness was a particular focus. Time-resolved measurements of wall static pressures and total pressure in the plume were correlated. A summary of the key findings is as follows:

• For nozzle pressure ratios that give rise to flow separation inside the nozzle, increasing the nozzle area ratio from 1 (straight nozzle) to 1.6 (convergent-divergent nozzle) results in a three-fold increase in the rms total pressure fluctuations near the nozzle exit. Spectra indicate that most of the instability energy is contained at low to moderate frequencies.

Proceedings of the International Conference on Aerospace Science and Technology 26 - 28 June 2008, Bangaloire, India

International Conference on Aerospace Science and Technology 26-28 June2008, Bangalore, India

• For the conditions of this study, the separation is asymmetric. • There are substantial correlations between the wall pressures caused by the flow separation and the total pressure inside the large separation zone. The frequency content of the total pressure fluctuation is similar to that of the separation motion.

• There is consistently better coherence between the total pressures in the separation zone. This suggests that the instability mechanism is due to an interaction between the expansion fans reflected from the smaller lambda foot with the shear layer of the larger separation zone.

GRAPHS Numerical and Experimental results were compared for the Mach 2 supersonic contour nozzle flow separation.

Fig. 2 shows the wall pressure distribution of Mach 2 supersonic contour nozzle for In-viscid case.

Fig. 3 shows the wall pressure distribution of Mach 2 supersonic contour nozzle for RNG model.

Proceedings of the International Conference on Aerospace Science and Technology 26 - 28 June 2008, Bangaloire, India

International Conference on Aerospace Science and Technology 26-28 June2008, Bangalore, India

Fig. 4 shows the wall pressure distribution of Mach 2 supersonic contour nozzle for K-ε model.

Fig. 5 shows the Experimental wall pressure distribution of Mach 2 supersonic contour nozzle.

REFERENCES 1. Dimitri Papamoschou and Johnson, A., "Unsteady Phenomena in Supersonic Nozzle Flow Separation" AIAA-2006-3360, presented at the AIAA 36th Fluid Dynamics Conference and Exhibit, June 4-11, 2006, San Francisco. 2. Xiao, Q., Tsai, H.M., and Papamoschou, D., "Numerical Investigation of Supersonic Nozzle Flow Separation," AIAA-2005-4640, presented at the 35th AIAA Fluid Dynamics Conference, June 6-9, 2005, Toronto, Ontario, Canada. 3. Romine, G. L., “Nozzle Flow Separation,” AIAA Journal, Vol. 36, No.9, pp.1618-1625, 1998. 4. Papamoschou, D. and Zill,.A., "Fundamental Investigation of Supersonic Nozzle Flow Separation," AIAA-2004-1111, presented at the AIAA 42nd Aerospace Sciences Meeting, January 2004, Reno, NV.