The application of the SCRAM2D code to investigating the flow fields that might be expected to occur in a representative Mach 10, two dimensional (ramp-compression) inlet is described. This Mach number allows the use of existing Navier-Stokes codes without the additional complexity of air chemistry and the associated large increases in computational time required to achieve numerical simulations of such flow fields. The CFD simulation consists of a two-dimensional inlet geometry that has an overall geometric turning of 36 degrees. The cowl for this inlet is assumed to be aligned with the oncoming freestream flow, turning the ramp flow field back parallel to the freestream, thus producing the overall 36 degree turning angle. The primary subject is the cowl shock wave-ramp boundary layer interaction effects. The results of numerical simulations carried out at the design Mach number of 10 and two other off-design Mach numbers, 7.2 and 5.0 are described

Rose, William C.

English

The flow in the transonic test facility was investigated using the three dimensional computational fluid dynamics techniques. The application of the full Navier-Stokes three dimensional code to the flow qualities in the contraction section of transonic wind tunnel is discussed. Initially, two dimensional solutions indicated the possibility for large secondary flow to exist as a result of the asymmetries involved in the contraction section as it is constructed. The results of a full three dimensional solution indicate that only minor pressure variations actually occur in the contraction section within any given cross flow plane. Further analysis of the three dimensional solution indicated that these slight lateral pressure gradients lead to negligible secondary flows, except within a small region in the corners within the boundary layer. On the basis of present solution, it would not be expected that any flow asymmetries and/or secondary flow present within contraction section are associated with the methods by which the contraction is implemented in its present configuration

NASA Technical Reports Server

Numerical investigations in three-dimensional internal flows

The present study is a preliminary investigation into the behavior of the flow within a 28 degree total geometric turning angle hypothetical Mach 10 inlet as calculated with the full three-dimensional Navier-Stokes equations. Comparison between the two-dimensional and three-dimensional solutions have been made. The overall compression is not significantly different between the two-dimensional and center plane three dimensional solutions. Approximately one-half to two-thirds of the inlet flow at the exit of the inlet behave nominally two-dimensionally. On the other hand, flow field non-uniformities in the three-dimensional solution indicate the potential significance of the sidewall boundary layer flows ingested into the inlet. The tailoring of the geometry at the inlet shoulder and on the cowl obtained in the two-dimensional parametric design study have also proved to be effective at controlling the boundary layer behavior in the three-dimensional code. The three-dimensional inlet solution remained started indicating that the two-dimensional design had a sufficient margin to allow for three-dimensional flow field effects. Although confidence is being gained in the use of SCRAM3D (three-dimensional full Navier-Stokes code) as applied to similar flow fields, the actual effects of the three-dimensional flow fields associated with sidewalls and wind tunnel installations can require verification with ground-based experiments

NUMERICAL INVESTIGATIONS IN THREE-DIMENSIONAL INTERNAL FLOWS
SEMI-ANNUAL STATUS REPORT
1 JANUARY THROUGH 30 JUNE 1991
,i , , ._
/V li
Prepared for:
NASA-AMES RESEARCH CENTER
MOFFETI' FIELD, CA 94035
UNDER NASA GRANT
NCC 2-507
ijr3 C ] _:_%
By:
WILLIAM C. ROSE
ENGINEERING RESEARCH AND DEVELOPMENT CENTER
UNIVERSITY OF NEVADA, RENO
RENO, NV 89557
https://ntrs.nasa.gov/search.jsp?R=19910016800 2020-03-19T16:50:24+00:00Z
I. BACKGROUND
In 1990 NASA initiated its Generic Hypersonics Research Program (GHP). The
general objectives of this research program are to develop technology background required
for aeronautical research in the hypersonic Mach number flow range. These research efforts
are to complement the National Aerospace Plane (NASP) program and are geared to the
development of experimental and computational fluid dynamics techniques. Previous
experience under the current research using the two-dimensional full Navier-Stokes code
(SCRAM2D) has indicated the desirability of producing highly contoured internal portions
of a hypothetical Mach 10 inlet. These results were presented in the previous progress
report. The two-dimensional code was used in a parametric sense to design the contours
for a specific Mach 10 hypersonic inlet. Flow conditions hypothesized to enter the inlet
were taken from the experimental conditions available in the NASA-Ames 3.5 foot
hypersonic wind tunnel. The 2D code has been used parametrically as a design tool because
of its reasonable results, ease of use and relatively short computer turn around time.
The effects that potential three-dimensional effects might have on such an inlet
designed with the two-dimensional code remain unknown. The purpose of the present
report is to describe the application of the three-dimensional full Navier-Stokes code
(SCRAM3D) to investigating the three dimensional flow fields that arise when swept
sidewalls are added to the two-dimensional compression lines obtained in the previous
portions of the present research effort.
° 1
II. INTRODUCTION
It is usually assumed for vehicles operating in the atmosphere above Mach numbers
of about 5 that the propulsion system will be highly integrated with the overall vehicle.
Because of the high temperature limitations of existing materials, control of the thick viscous
boundary layer entering the propulsion path as a result of the highly integrated forebody is
expected to have limited practical application at the high cruise Mach numbers. Thus it is
desirable to develop an understanding of the flow fields expected to occur in such an
integrated inlet system. At the high Mach numbers handling the thick boundary layers may
be relatively easy due to their higher momentum content. However, as the Mach numbers
decrease it is known that even the turbulent boundary layer will become more liable to
separation and potential implications of inlet operability arise. For these cases, boundary
layer bleed (below Mach numbers of about 5) may be required as part of the inlet operating
procedure. However, for purposes of the present investigation, we concentrate on the cruise
Mach number (assumed to be 10), and no boundary layer bleed is considered here.
In the previous progress report, various designs were arrived at using the SCRAM2D
code. All of the designs have a geometry that is similar to the one shown in Figure 1 which
is repeated from the previous report. The specific geometry shown in Figure 1 is noted as
Mod. 23 and has a 10 degree compression forebody simulation section followed by two 4
degree ramps that nominally produce a coalescent shock wave system just ahead of the cowl
lip. The specific case shown in Figure 1 also has a straight cowl with a slightly rounded
ramp shoulder and a straight, constant area throat. This geometry produces a 36 degree
total geometric turning angle. Another version studied extensively was one having a 28
2
degree total geometric turning angle, which was produced by having only one 4 degree
compression on the ramp surface. Flow fields calculated from the 2D code for the inlet
shown in Figure 1 are shown in Figure 2, and indicate the difficulties that occur in obtaining
a smooth pressure distribution in the throat without contouring of the internal flow surfaces.
Figure 3 shows the flow field that results when both cowl contouring and ramp contouring
are used. It was shown in the previous study that contours such as these tend to produce
more uniform flow fields in shorter throat (isolator) regions than without contouring. These
results are repeated from the previous progress report. New efforts are discussed below.
Similar improved flow field behavior was found to result with the contoured ramp
and cowl surfaces for the 28 degree geometric turning angle (Mod. 26G) as shown in Figure
4. The two-dimensional contour used in Figure 4 is the one chosen for the three-
dimensional study, since the pressure rises are lower than the 36 degree case shown in
Figure 3 and are more representative of those required for a realistic flight environment.
The design Mach number of 10 was used also in the 3D study.
The Mach 10 design value was chosen to be a leap in the required technology beyond
that developed in the NASP "Mach 5" inlet study. The Mach number of 10, on the other
hand, is still low enough so that real gas/air chemistry issues such as dissociation and
ionization are not expected to be the dominant issue in establishing the performance of the
inlet either in ground-based simulations or in flight along NASP-like trajectories
(approximately 2,000 psf dynamic pressure). These facts allow us to continue using the full
Navier-Stokes codes, both SCRAM2D and SCRAM3D without the additional complexity
3
of air chemistry and the associated increases in computational time required to achieve
numerical simulations of these flow fields.
HI. RESULTS AND DISCUSSION
The two-dimensional compression contours shown for the solution in Figure 4 were
applied to a three-dimensional grid having 301 points in the streamwise direction by 61
points across the two-dimensional compression direction by 31 points in the spanwise
direction. The inlet was assumed to be symmetrical and have a half-width of 0.20 meters.
The sidewalls of the grid were assumed to originate at the leading edge of the 10 degree
compression ramp and terminate at the lip of the cowl. Figure 5 shows the perspective view
of the inlet geometry with the ramp at the top, the center plane of the inlet shown to the
left side with the sidewall to the right and the cowl in the lower portion of the figure. Flow
is assumed to be from right to left in this figure. The cross-flow planes shown contain the
Mach number contours from the three-dimensional solution obtained in the present study.
Figure 6 adds the Mach number contours on the center plane of the inlet.
For modular installations, adjacent inlets share sidewalls. On the other hand, for
purposes of wind tunnel testing, usually an isolated inlet is considered. For purposes of
computational consideration, these two situations distinguish themselves primarily by either
having an outboard ramp, that is a ramp beyond the sidewall, or not having an outboard
ramp. In the present study, these two cases were examined with the multi-block SCRAM3D
code for treating both the flow outboard of the sidewall as well as below the edge of the
sidewall. Figure 7 shows the Mach number contours in a cross flow plane just ahead of the
4
cowl lip for the casehaving an outboard ramp similar to the multi-module installation.
Little flow migrates across the sidewall, since the sidewall plane acts like a plane of
symmetry. A large sidewall vortical feature (characteristicof ramp compressioninlets with
sidewalls) is seen(2). There is lesstendencyof the flow to spill over the sidewall (1). Due
to the large displacementeffect of the vortical feature on the flow field, both the initial and
secondramp shockwavesare bowed in the region of the sidewall. The flow field behavior
shown in this figure is in contrast to that shown in Figure 8.
In Figure 8, the solution is shown at the same streamwise station as in Figure 7,
however, for this case the outboard ramp has been removed so that the three-dimensional
simulation is similar to that expected for an isolated inlet model as tested in a wind tunnel.
In this case, the lower pressure associated with the freestream of the wind tunnel is allowed
to exist on the outer boundaries of the solution. With this lower pressure outboard of the
sidewalls (1), large spillage occurs over the sidewalls ahead of the cowl lip (2). This spillage
tends to reduce (but not eliminate) the size of the vortical feature that is ultimately ingested
into the internal portion of the inlet. This also results in the ramp shock wave having a
slightly shallower angle along the sidewall, as can be seen in Figure 8.
In previous 3D studies, the sidewall flow fields have been shown to cause large
perturbations to the internal flow fields, while the center portions of such inlets remain
nominally two-dimensional. The differences between the two-dimensional and three-
dimensional surface pressure distributions (from the center plane of the inlet) are shown in
Figure 9. Some minor differences occur; however, the overall compression achieved along
5
the center line of the 3D solution is quite similar to that obtained with the 2D solution.
Figure 10 shows a cross flow plane mass flow profile obtained at the exit of the inlet
(around 2.0 meters). Here, the nominally two-dimensional portion of the inlet flow is seen
dearly, while the flow near the sidewall shows the remnants of the sidewall vortex as has
been acted on by the cowl shock wave and subsequent internal compressions.
The effect of three-dimensional flow fields on inlet performance is of interest. For
purposes of comparison, Table 1 shows the mass-averaged total pressure recovery for the
2D and 3D calculations at three streamwise positions: forward of the cowl shock, just aft
of the cowl shock, and near the exit of the inlet (denoted by the term "throat"). The
magnitudes of the total pressure recoveries are not as significant as the differences between
the 2D and 3D calculations as one goes through the inlet. For the overall recovery at the
exit of the inlet, the ratio of the two-dimensional to three-dimensional calculated values is
1.31. This says that by doing 2D calculations alone one is likely to overestimate the inlet's
performance when compared with that calculated with the 3D code (having the losses
associated with the sidewall vortical flow and its interaction throughout the internal flow
portion of the inlet).
21)
3D
FORWARD OF AFT OF 2D/3D
COWL SHOCK COWL SHOCK THROAT RATIQ
0.370 0.108 0.038
0.329 0.082 0.029
Table 1. Mass Averaged Total Pressure Recovery
1.31
IV. CONCLUDING REMARKS
The present study is a preliminary investigation into the behavior of the flow within
a 28 degree total geometric turning angle hypothetical Mach 10 inlet as calculated with the
full three-dimensional Navier-Stokes equations. Comparison between the 2D and 3D
solutions have been made. The overall compression is not significantly different between
the 2D and center plane 3D solutions. Approximately one-half to two-thirds of the inlet
flow at the exit of the inlet behave nominally two-dimensionally. On the other hand, flow
field non-uniformities in the three-dimensional solution indicate the potential significance
of the sidewall boundary layer flows ingested into the inlet.
The tailoring of the geometry at the inlet shoulder and on the cowl obtained in the
2D parametric design study have also proved to be effective at controlling the boundary
layer behavior in the 3D code. The 3D inlet solution remained started indicating that the
2D design had a sufficient margin to allow for three-dimensional flow field effects.
Although confidence is being gained in the use of SCRAM3D as applied to similar flow
fields, the actual effects of the three-dimensional flow fields associated with sidewalls and
wind tunnel installations can require verification with ground-based experiments.
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(b) Cowl surface pressure distribution.
k180.0
Normalized Pressure on Ramp Surface
M10 Inlet Mod 26G M=10
2D solution
3D solution without outboard ramp
150.0
120.0
90.0
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pressures.
of two.dimensional and three-dimensional (center plane)
(a) Ramp surface pressure distribution.
mini


Numerical investigations in three-dimensional internal flows

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