cfd simulation of s-duct test case using overset grids conference...american institute of...

American Institute of Aeronautics and Astronautics

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CFD Simulation of S-Duct Test Case Using Overset Grids

L. M. Gea1

The Boeing Company, Huntington Beach, CA 92647, USA

Steve Nyugen2

California State University at Long Beach, Long Beach, CA 90840, USA

The primary objective of this paper is to assess the capability of overset computational

fluid dynamics (CFD) code to simulate the steady-state aerodynamics of compact offset

intake diffusers. CFD simulations are conducted for Serpentine S-duct configuration, and

the predicted results are compared with measured wind tunnel data including boundary

layer profiles, surface static pressures and total pressure in the region of interest. Numerical

aspects such as y+, turbulence model, and free stream Mach numbers are also studied to

enhance the code applicability. This paper is a summary of the work conducted for 1st

Propulsion Aerodynamic Workshop (PAW01).

I. Introduction

ue to the rapid advancement for the modern propulsion airframe integration, the demand for a reliable CFD capability to predict propulsion aerodynamic has increased significantly in recent years. In order to encourage

the collaboration between industry and universities and promote the state-of-art simulation techniques, 1st

Propulsion Aerodynamic Workshop (PAW01), organized by AIAA Air Breathing Propulsion System Integration

Technical Committee (ABPSI), was held during the 2012 Joint Propulsion Conference in Atlanta Georgia.

One of the test cases selected by the workshop committee is the Serpentine diffuser S-duct configuration

provided by Anne-Laure Delot of ONERA1. The experimental model (Figure 1) of this configuration has been

extensively tested at the ONERA wind tunnel. The measured boundary layer profiles were provided for all the

participants to gauge their CFD results. Surface static pressure as well as total pressure at given locations were

required by PAW01 committee to evaluate the overall performance of each individual CFD code.

OVERFLOW2, a NASA developed general purpose Navier-Stokes code using Chimera overset technique3, has

been widely adopted by the CFD community throughout the industry4-9. The emphasis of using OVERFLOW for the

propulsion aerodynamic application within the Boeing Company in recent years highlights the success of this versatile code. However, majority experience accumulated so far was primarily focused on external aerodynamics.

There is an urgent need to expand the applications to internal flow aerodynamic predictions. The S-duct test case for

PAW01 provides a good opportunity to extend the knowledge base.

A careful designed overset grid system was generated to best simulate the wind tunnel test model, with the

consideration of including the model geometry needs as well as the flow physic features. By comparing the

predicted boundary layer profiles and surface static pressures, several numerical aspects, such as y+, turbulence

model, and free stream Mach, were studied to establish guidelines for the subsequent CFD simulation. The CFD

predicted boundary layer profiles, Mach contours, and total pressure contours for a specific mass flow rate are

compared with test data.

II. Experimental Model And Test Cases

The experimental model (Figure 1) was manufactured and tested at the ONERA R4MA wind tunnel facility

based on a smaller scale model developed by Harloff et al. at NASA Lewis Research Center in the 1990’s10. The

model is composed of a bell mouth, a constant diameter pipe, and S-shaped duct. The area ratio of the S-duct (ratio between the outlet and inlet section) is equal to 1.52 with the inlet diameter (D1) of 133.15 mm and outlet diameter

(D2) of 164 mm. The outlet section is connected to the Aerodynamic Interface Plane (AIP), located at s/D1=5.587,

with 40 Kulite pressure probe on it. S is the distance measured along the S-duct centerline from x=0.

1 Associate Technical Fellow, The Boeing Company, [email protected], AIAA Senior Member. 2 Graduate Student, Mechanical/Aerospace Engineering Department, CSULB, [email protected].

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The CAD model, shown in Figure 2, illustrates the test case geometry with color coded to distinguish the

purpose of each segment. The length unit and angles shown in the figure is in millimeters and degrees respectively.

The centerline inside the red colored S-duct region originates from two equal radius 30o circular arcs tangent to each

other. The ratio of cross section diameter (D) to the inlet diameter along the S-duct and detailed geometric definition

is given by Harloff et al. The experimental model contains the bell mouth inlet (blue) up to and including the

constant diameter pipe (green) following the S-duct (red). The constant diameter region allows the incoming flow to become fully developed to a desired Mach number imitating free stream fight conditions. Boundary layer probes

were situated at x=-76.5 (s/D1=-0.575) to record the fully developed pressure and velocity profiles before air enters

the S-duct at x=0 (s/D1=0). Once in the S-duct, wall pressure sensors were utilized to measure the ratio of surface

stagnation pressure to free stream total pressure. Axial pressure taps were positioned at the polar angles (ϕ) of 0o,

90o and 180o, and circumferential pressure taps were placed at s/D1=2, 3, and 4 as shown in Figure 3.

Two test cases with different mass flow rates were performed on the experimental model by Delot et. Al.

Mass flow rate of 2.427 kg/s in the first case resulted in a Mach number of 0.3549 at AIP and 0.6 at inlet. Reducing

the mass flow rate to 1.356 kg/s for second case resulted in Mach number of 0.1819 at AIP and 0.4 at inlet. These

two cases were selected by the PAW01 committee as the CFD test cases, and measured boundary layer profiles were

provided for participants to validate their CFD simulation. Subsequently, CFD predicted surface pressure as well as

Mach and total pressure contours at AIP were requested by PAW01 from each participant. The data were collected

before the workshop and overall evaluation and comparison conducted by the committee were presented during the workshop.

Figure 1. Serpentine diffuser wind tunnel model


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Figure 2. CAD model with dimensions and CFD model coordinates

Figure 3. Axial and circumferential pressure taps locations on half model


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III. Computatioal Model And Flow Solver

An overset grid approach was used for this CFD task. Six structured overlapped grids, shown in Figure 4, were

generated separately with each grid covering a portion of the flow field surrounding the S-duct configuration. Due to

the symmetric nature, only half model was considered for current study.

Grid number 1 is an O-type viscous grid which grows out from the solid wall surface for about 40 mm. Grid

number 2 is an H-type inviscid grid covering the center of the S-duct from the inlet entrance to AIP. This particular

grid avoids the often troublesome axis treatment for circular shape of internal flow simulation. Grid number 3 is

also an O-type viscous grid which grows out from the solid surface of the spinner for about 50mm. Grid number 4

is an O-type viscous grid caps the nose of the spinner. Grid number 5 is an O-type inviscid gird covering the

exterior portion outside of the bell mouth. A separate gird (number 6) was generated to prevent the interpolation associated with grid overlap at the exit plane, where mass flow needs to be computer during the run to monitor the

convergence. In total, there are about 11 million grid points for the entire grid system. The interpolation information

between the six overlapped grids were obtained by using Pegasus5, a NASA developed domain connectivity code.

OVERFLOW, a NASA developed general purpose Reynolds-averaged Navier-Stokes (RANS) code, with Chimera

overset scheme incooperated, was employed for this study. OVERFLOW code has been extensively validated and

widely used within the Boeing company for many airplane development programs. The success in the propulsion

airframe integration (PAI) analysis is one of the highlights. However, majority of the experience accumulated so far

has been for the external aerodynamics. Relative limited experience for internal flow simulation has been

accumulated. In order to meet the challenge of growing demand for variety of CFD applications, it is important to

understand the OVERFLOW applicability for internal flow simulation. The S-duct test case for PAW01 offered a

great opportunity to explore this area.

Figure 4. Overset grid system


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IV. Numerical Studies

In order to ensure the quality for the CFD results, three separate studies were conducted to establish the best

practice for the current CFD simulation. Although the studies here were focused on the current test configuration,

the lessons learned and issues addressed can be general guidelines and applied to other applications.

A. Grid Convergence

One important aerodynamic data provide by PAW01 to gauge the CFD predicting accuracy are the boundary

layer profiles measured from a point located in the constant diameter region downstream of the inlet (x=-76.58).

Another date also obtained from the wind tunnel test was the surface static pressure in the curving S-duct region

between x=-200 and 800. This data were later used by the committee to evaluate the performance between different

CFD codes. Since both quantities are of boundary layer related, a grid system with adequate y+ is essential. Three different grid systems with y+ of 0.1, 1.0, and 10.0 were generated. The CFD predicted boundary layer profiles,

along with test data, and the static pressures along the S-duct surface in three different azimuth angles were shown

in Figure 5a to study the grid convergence. Ignore the correlation with test data for now, it is clearly shown that the

predicted boundary layer profiles converge once y+ becomes smaller than 1. Similar trend was also observed from

the surface static pressure plot. It is interesting to point out that the surface pressure becomes identical downstream

of x~250. Further studies show that flow starts to separate from about x=250, and the wall pressure becomes

constant in the separated region, therefore in-sensitive to y+. It is then determined to use y+ of 1 for the rest of the

study.

Figure 5a. Grid convergence study, y+ = 0.1, 1.0, and 10.0


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B. Turbulence Model

Two most widely used turbulence model, two equation SST model11 and one equation Spalart-Allmaras (SA)

model12, were experimented for the current test case. The OVERFLOW predicted boundary layer profiles and

surface static pressures using both models were shown in Figure 5b. It is shown that the boundary layer profiles

predicted by both models deviate from the test data, and no clear advantage for either model can be declared. For the

surface static pressures, similar results were predicted by both models upstream of x~200, larger deviation can be observed in the curving region of the S-duct. This is expected as different turbulence model behaves differently in

the separated region. SST model was chosen for the rest of the study.

Figure 5b. Turbulence model study, SST vs. SA models

C. Free Stream Mach Number

In order to simulate the wind tunnel environment, a zero Mach number condition is often required for the far

field free stream condition. However, OVERFLOW as a compressible code prohibits the zero Mach condition. A

very small Mach number is usually imposed for many external CFD applications. The intent here is to study the free

stream Mach number effect to internal flow simulation. OVERFLOW predicted results using three different free

stream Mach numbers, 0.005, 0.01, and 0.02, were compared in Figure 5c. At azimuth angles of 0o and 180o, the predicted pressures are almost identical near the wall but the difference becomes pronouncing moving away from

the wall. The profile for Mach=0.02 case shows largest deviation. No difference can be observed at ϕ= 90o. The

predicted pressures along the S-duct wall surface at all three azimuth angles are almost identical for all three Mach

numbers. It is demonstrated here that varying the free stream Mach number has little effect to the surface quantities,

but not necessary true for the field quantities. Based on the study, free stream Mach number of 0.01 was chosen for

the study. However, it is recommended that more in depth studies are needed to further understand the impact of

using different small free stream Mach numbers for internal flow simulation.


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Figure 5c. Free stream Mach number study, Mach=0.005, 0.01, and 0.02

V. Results and Discussions

Two test cases, one standard and one optional, were required by PAW01 for CFD simulation. Both results were

submitted for the workshop, however, only the standard test case will be presented in this paper.

The standard test case has a mass flow rate of 2.47 kg/s resulting in a Mach number of 0.5849 at S-duct inlet.

The equivalent Reynolds number is about 230 per millimeter based on the far field tunnel condition. OVERFLOW

predicted Mach and pressure ratio contours on the symmetric plane and AIP station are shown on Figure 6.

Streamlines on the symmetric plane are also included to facilitate the flow analysis. After entering the inlet, a

uniform flow was developed in the straight constant diameter region. Based on CFD prediction, Mach number

reaches about 0.6 which agrees well with the measured Mach number from the test. Moving further downstream in

the curved portion of the S-duct, the flow decelerates due to the geometric shape as well as the blocking effect from

the spinner located downstream of the AIP. Adverse pressure gradient triggers the flow to separate, starting from about x=250 on the ϕ=180o side, and a low velocity region can be observed extending pass the AIP station into the

spinner region. The flow on the ϕ=0o side remains attached throughout the S-duct. The Mach and pressure contours

on the AIP station show the extent of the separation core inside the duct.

OVERFLOW predicted boundary layer profiles are compared with test data in Figure 7. Although in general

compared quite well, some distinct discrepancy can be found in all three azimuth angles. During the workshop, it

was found that similar discrepancy was reported by almost all the participants. It is suspected that there may be

some inconsistencies between the experimental model and the geometry that was actually modeled in CFD. CFD

predicted pressure along the S-duct surface is compared with measured data in Figure 8. Large discrepancy was

observed for entire ϕ=180o cut except within the separated region (between x=250 and 400), where pressure is

nearly constant. For ϕ=0o and 90o, good agreement can be found between CFD and test data except the region

upstream of x=300.


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One of the most important aspects for the inlet aerodynamic simulation is to predict the pressure recovery and

distortion indices. During the wind tunnel test, AIP measuring rake, consists of 8 arms, of 5 pressure probes each,

was used to measure the pressure data for CFD validation. In order to compare the CFD results with these field

quantities obtained from the test, two half model rake grids with different probes density, one wind tunnel used rake

like 5X5 and the other 13X5 with more probes, were generated. And the solution on the rake grids were interpolated

from the CFD results obtained using the original CFD gird. Mach and pressure contours shown on the original grid and two rake grids are compared in Figure 9. The area averaged pressure, often used when determining recovery and

distortion indices, is also computed and shown with the corresponding grid. It is found that this parameter is not

sensitive to the probe density of the rake grid; therefore, the coarse 8X5 (full model) grid used in the test is probably

adequate for recovery and distortion calculations.

Finally, the CFD predicted pressure contours at AIP are compared with test data in Figure 10. Good correlation

was found except within the separated region where high gradient occurs. It is important to point out that the

experimental color contours shown on the left hand side in 13X5 grid (half model) were interpolated from the data

measured from a coarser 8X5 rake in the test. And the CFD predicted contours shown on the right hand side were

post processed in a similar procedure. Data for a coarse 8X5 rake grid was first generated by interpolating solution

from the original CFD results, the solution is then interpolated onto the 13X5 rake grid.

Figure 6. Mach and pressure contours on symmetric plane and AIP station


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Figure 7. OVERFLOW predicted boundary layer profile vs. test data

Figure 8. OVERFLOW predicted surface pressure vs. test data


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Figure 9. CFD predicted Mach and pressure contours on rake grids

Figure 10. CFD vs. experimental pressure contours at AIP


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VI. Conclusion

The current analyses have demonstrated the capability of an overset CFD code to properly simulate the internal

flow for the serpentine diffuser S-duct configuration. On the practical side, the boundary layer profiles, surface

static pressures along the S-duct, and the Mach and pressure counter at AIP are adequately predicted and in good

correlation with wind tunnel test data. The wind tunnel measured boundary layer profiles deviate from all the CFD

predicted results suggests the need to revisit the consistence between experimental and CFD geometries. Previously

common practice of using Mach number of 0.01 for zero Mach external flow applications needs to be carefully

validated for internal flow applications.

Acknowledgments

The first author would like to thank PAW01 for selecting this work as an invited paper.

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cfd simulation of s-duct test case using overset grids conference...american institute of...

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