AIAA 2011-3340
29th AIAA Applied Aerodynamics Conference
27 - 30 June 2011, Honolulu, Hawaii
Turbulence Model Assessment for Hot Plumes
Andrea J. Shestopalov,
∗
Robert E. Childs,
I.
†
John E. Melton,
‡
Introduction
This work was motivated by the need to accurately characterize the aerodynamic behavior of the Orion
Launch Abort Vehicle (LAV). The abort system is powered by solid rocket motors, and the vehicle’s aerodynamics involves complex interactions between multiple rocket plumes and the freestream. The LAV’s
aerodynamic behavior is predicted using a combination of wind tunnel testing and CFD analyses. The goal
is to rely on wind tunnel data to the extent possible. However, numerous issues, like the difficulties of using
solid rocket motors in a wind tunnel, make realistic and accurate experiments difficult to perform. CFD
is used to reach the conditions which are difficult or impossible to achieve experimentally. CFD also has
limitations, and one of the most significant is turbulence modeling. This work addresses turbulence modeling
for rocket motor plumes, and particularly, the effects of high-temperature jets. This is a significant concern
because the majority of wind tunnel tests use ambient temperature compressed air for plumes, while the
stagnation temperature of flight plumes is an order of magnitude higher. Issues of rocket plume chemistry
are beyond the scope of this study.
The scope of this work is both to validate that an available CFD model can correctly simulate the
experimental results, and to select an appropriate CFD model for future hot plume predictions. The CFD
model selection and validation involves comparisons of several turbulence model results to data from a
single underexpanded supersonic jet inclined to create a 25◦ or 40◦ angle between the jet and a M = 0.3
freestream. Jets with stagnation temperatures of 530◦ R and 1350◦ R were studied. The work focuses on the
SST turbulence model, but results from other 1- and 2-equation models are also given. A compressibility
correction which reduces the spreading rate of high-speed mixing layers, and a temperature correction (AbolHamid) which increases the spreading rate of hot jets were studied. Comparisons of CFD and experiment are
given for PIV data of the hot-jet velocity field, and for surface pressures on an LAV-like capsule for hot and
cold jets. Water-vapor condensation precluded PIV measurements in the cold-jet cases. The experimental
work has been given the designation “85-AA” within the Orion Crew Exploration Vehicle (CEV) Aerosciences
Project and hereafter will be referred to as 85-AA.
II.
Experiment
The 85-AA experiment was performed at NASA Glenn Research Center in the AeroAcoustic Propulsion
Laboratory (AAPL). The AAPL is an acoustic test facility housed in a 130 ft. diameter acoustically lined
dome. The experiment was run in the Nozzle Aeroacoustic Test Rig (NATR) which is a 53 inch diameter open
wind tunnel capable of freestream Mach numbers up to M = 0.3, and this maximum value was used for the
test. High pressure air with optional heating for the test jet was provided via an engine simulator within the
NATR, referred to as the High Flow Jet Exit Rig (HFJER). The hot jet was vitiated air resulting from natural
gas combustion with a ratio of specific heats γ ∼ 1.32. The HFJER was intended to provide high flow rates at
the moderate pressure and temperature typical of the jets from modern aircraft engines. The small energetic
jet required for this test was at the limit of the jet supply system. The hot jet was set to a total temperature
of ∼ 1350◦ R, while the cold jet was ∼ 530◦ R. The nozzle pressure ratio (N P R = Pjet stag. /Pambient ) was
28.5 for hot and cold jets; the mass flow rate was approximately 1 lbm/sec.
The test involved a single jet at three different angles with respect to the freestream and an optional 4%
scale model capsule. The experimental test rig is shown in Figs. 1 and 2. The jet and capsule had to be
separated by free air to enable good access for the PIV laser sheet without the possibility of reflected light
∗ Research
Scientist, STC (currently Project Manager, Exa Corp.)
Scientist, STC
‡ NASA Ames Research Center
† Research
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contaminating the measurements. Three nozzle configurations were used, with angles of 0◦ , 25◦ , and 40◦
from the freejet axis. The 25◦ and 40◦ nozzle cases simulate 0◦ and −15◦ angles of attack, and were used
with the capsule tilted 0◦ and −15◦ , respectively. The zero degree nozzle case was not intended for use with
the capsule. The capsule’s vertical location was chosen for both non-zero nozzle angles to cause the plume
to graze the capsule, with the goal of producing a strong interaction between the plume and capsule.
Measurements included capsule surface pressures and field velocities in the hot-jet cases. The capsule had
63 static pressure taps arranged as shown in Figs. 3 and 4. Due to water condensation in the cold-jet cases,
the surface pressure data on the capsule provide the only means of making a direct comparison of turbulence
modeling accuracy between hot and cold jets. Velocity field data were acquired with Stereo Particle Image
Velocimetry (SPIV).1 The SPIV system used two high resolution cameras (4008 x 2672 pixels) to provide a
600 mm square field of view. The jet flow was seeded with alumina particles, and the wind tunnel flow was
seeded with propylene glycol. PIV data were acquired in crossflow planes at the stations noted in Section
3.3.2 and a streamwise-vertical plane on the centerplane of the plume. A sequence of 400 velocity vector
maps were acquired at each measurement station and averaged to provide first and second order statistics
over each plane.
The experimental results for the capsule surface pressures showed both asymmetry and variance from
repeat runs. Uncertainty bars were calculated for the capsule pressure data using the following equation,
q
(1)
UCP = [(σsym )2 + (σrepeat )2 ],
where σsym and σrepeat are the estimated variances due to symmetry and repeatability, respectively, calculated with a range analysis of symmetric or repeated runs.
Figure 1: 85AA Experiment in the NATR
Figure 2: CAD Model of 85AA Geometry
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Figure 3: Pressure Taps on Capsule
III.
III.A.
Figure 4: Pressure Taps on Capsule
CFD Model of 85-AA
Geometry
The simplified geometry used for the CFD model includes the NATR shroud, HFJER strut, nozzle, and
capsule. The coordinate system of the CFD geometry is such that the X direction is downstream, the Y
direction is cross-stream, and the Z direction is verticle. The coordinate system of the experiment, however,
is such that the Z direction is downstream and the X direction is cross-stream. The CFD results were then
rotated to correspond with the experimental coordinate system for all comparisons.
Not included in the model are the capsule-support sting and the so-called “mini-strut” which supports
the jet nozzle. The CFD entire geometry is shown in Fig. 5, with the nozzle and capsule isolated in Figs. 6
and 7 to highlight the 25◦ and 40◦ nozzle configurations, respectively. For the turbulence modeling studies,
the jet inflow is defined by boundary conditions just upstream of the nozzle throat, neglecting the interior
ducting of the nozzle sting, strut and the HFJER plenum. A CFD solution of the flow in this internal
plumbing, shown in Section 3.3.3 suggests that details of the internal flow have a modest impact on the CFD
results.
Figure 5: CFD Geometry
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Figure 6: CFD Geometry, 25 degree nozzle
III.B.
Figure 7: CFD Geometry, 40 degree nozzle
CFD Cases
At the onset of this study, our first goal was to demonstrate the suitability of the modeling approach. We
will begin with an assessment of each of the following assumptions:
• Grid Converged Solution
• Using air to model plume rather than multiple species
• Neglecting interior ducting
• Neglecting additional stings and supports
Following the validation of our primary assumptions, we will present the results of a study of the different
turbulence models in matching the PIV plume data from the experiment for the hot jet case. We will then
use the PIV data to investigate the performance of the SST model with and without corrections on modeling
the hot jet plumes, and we will conclude with the comparison of capsule pressures to investigate the disparity
in accuracy of modeling cold and hot jets.
III.B.1.
Grid Convergence
An initial grid convergence study was done with the standard Mentor SST2 turbulence model that indicated
a grid-converged solution, with the baseline plume grid, based on the plume core velocities. However, the
plumes generated using the compressibility-corrected (CC) SST model gave thinner shear layers and had
core velocities that decayed less rapidly, in comparison to the standard SST model. Thus, an additional
SST-CC grid convergence study was performed. Both grid refinement studies addressed the plume grids by
modifying the plume x factor and plume r factor variables that scaled the default grid spacing in the (jet)
streamwise and radial directions, respectively. The values for these variables, as well as the resulting grid
sizes, are shown in Table 1. Figure 8 shows the plume centerline velocities for each of the grids using the
SST-CC model. The baseline grid, Grid #1, shows a slight difference in peak velocities compared to the
other two grids, but results from Grid #2 to Grid #3 are very similar. Thus, Grid #2 provides well-resolved
solutions for all turbulence models, and it was chosen for all simulations in this work.
Grid
Grid #1: baseline
Grid #2: Refine 1.5
Grid #3: Refine 2.0
plume x factor
plume r factor
grid size
0.667
0.5
0.333
0.4
0.3
0.2
25.4 million grid points
28.8 million grid points
38.1 million grid points
Table 1: Grids used in refinement study
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Figure 8: Centerline Velocities of Grids for Convergence Study
III.B.2.
Hot Jet Gas Modeling
The hot jet is vitiated air with a ratio of specific heats of γ = 1.32, and the significance of modeling the jet
as hot air was quantified. Overflow3 has the option to approximate multiple different gases with a ‘frozen
chemistry’ model. Using the jet-only case at 25 degrees, we compared the simulated plume using air and
then using a second species. Figure 9 shows the experimental PIV slice locations for the 25 degree nozzle.
Figure 10 shows cross-sectional slices of the plume at each PIV slice location. The slices in Fig. 10 show the
velocity magnitude of the experiment as a filled color contour, and the velocity magnitude of the Overflow
simulation as an overlaying black line contour. The left hand columns show the experiment and the Overflow
simulation with air, while the right hand columns show the experiment and the Overflow simulation using
multiple species. The contour levels for each slice are identical, showing a slight increase in the core velocity
for the multiple species simulation, but a very small change in the shape of the plume. The simulation
contours in Fig. 10 were all obtained using the standard SST model with no corrections. Fig. 11, generated
in the work by N. Gross,4 shows the velocity profiles given by the SST (ICC=0, ITC=0) and SST-CTC
(ICC=1, ITC=1) models on a vertical traverse through the center of the jet at X/D = 18 . The difference
in core velocity is more prominent with the addition of the compressibility and temperature corrections, but
the differences are still small in absolute terms, with respect to the variation due to turbulence model choice.
This demonstrates the validity of using air to model the jet exhaust for this experiment.
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Figure 9: PIV Slice Locations, 25 deg. nozzle
SST, air
SST, species
SST, air
slice 1
slice 4
slice 2
slice 5
slice 3
SST, species
slice 6
Figure 10: PIV slices, 25 deg. nozzle, hot jet, multiple species v. air
Figure 11: Velocity Profile, multiple species v. air
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III.B.3.
Nozzle Interior Ducting
The CFD nozzle flow was modeled using uniform ideal inflow boundary conditions in the plenum just
upstream of the nozzle throat, and the sensitivity of the CFD results to this assumption was evaluated. The
flow-path leading to the nozzle throat was sized to have flow at M ∼ 0.3, based on idealized one-dimensional
flow. This conformed to the best practices used in experiments at the AAPL facility. Comparisons of
experiment and CFD displayed a difference in plume shear layer thickness at the first measurement slice
downstream of the nozzle exit, and was outside the range of all models tested. The simulated plume shear
layer was much thinner than the shear layer in the experiment, calling into question the validity of the
assumption to neglect the rig’s interior ducting. This behavior is shown in Fig. 12. The color contours are
the experimental PIV measurements, and the overlaid black contour lines are the CFD runs. Both the SST
and SST-CC models give a thinner shear layer than the measured experimental results.
Many of the possible causes of this difference were evaluated by testing alternate nozzle inflow conditions
in the CFD. Simulations modeling the full internal flow path from the HEFJR to the nozzle were performed
using the Star-CCM flow solver, to quantify these potential effects. A cross section from the Star-CCM
simulation of the interior flow is shown in Fig. 13 as total pressure contours. Although the flow goes through
a number of turns, the velocity magnitude slice just aft of the nozzle exit (slice 0), shown in Fig. 14, does not
exhibit a thick shear layer as seen in the experimental color contours in Fig. 12. Using the velocity profile from
the Star-CCM simulation of the full interior flow path as a boundary condition for the Overflow calculations
yielded a negligible change in the external plume shear layer. The prevailing belief is that most of the energy
in the nozzle inflow is in the pressure, so even moderate non-uniformity in the velocity disappears as the
flow goes through the throat. Increasing the turbulence kinetic energy and eddy viscosity at the plenum
inflow boundary also had negligible effect on plume development, but this was expected, as turbulence in the
nozzle’s core flow decays very rapidly. This set of results suggests that no form of modification to conventional
RANS modeling of this flow can bridge the differences between CFD and experiment. In a review of LES
of jet aeroacoustics, Bodony and Lele5 note that nozzle boundary layer thickness and unsteadiness greatly
affect the predicted mean flow and acoustics in the resulting jets. It is worth noting that the nozzle interior
walls were produced by drilling and electrical discharge machining (EDM) and had small scale roughness.
This leads to the conclusion that there is either some aspect of the internal geometry that we cannot model
(i.e. interior grooves due to manufacturing processes), or with the available CFD methods we are not capable
of correctly capturing all the complexities of the flow. With the information currently available, the inclusion
of the internal flow in the computations does not affect the plume shape just aft of the nozzle, validating
our assumption to neglect the ducting.
Figure 12: Overflow SST (right) and SST-CC (left) velocity contours v. Experiment at nozzle exit
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Figure 13: Internal ducting modeling with Star-CCM
Figure 14: Velocity contours at nozzle exit with Star CCM, SST model
III.B.4.
Additional Stings and Supports
The validity of using a simplified geometry for CFD was tested. The simplified CFD geometry used for most
simulations includes the HEFJR, capsule, and nozzle body, as shown in Figs. 6 and 7, but not the nozzle
mini-strut, HEFJR sting, or capsule support of the complete geometry. This simplified geometry was used
initially, before the complete geometry had been included in the grid generation process. When the full
geometry was completed, it was tested and found to have an effect on the results. The results from full and
simplified geometries are compared here using the 40 degree nozzle configuration with capsule, both with a
hot jet and no jet.
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The capsule slices where pressure measurements were taken are shown in Fig. 15. Circumferential slices
of the pressure taps at the aft end of the capsule (Cut 4 in Fig 15) are shown in Figs. 16 and 17 for a hot
jet case with and without the HEFJR sting and capsule supports. Figure 16 shows the effect of adding the
sting with the standard SST model, while Fig. 17 shows the SST-CC model. The distance between the
CFD curves and experiment near Φ = 0 and Φ = +/ − 180 goes down considerably when the HEFJR sting
and capsule support are included. The addition of the HEFJR sting and capsule support move the pressure
closer to the experimental data with the jet on. The assumption that neglecting the stings and supports
would not effect the flow was incorrect. Thus the geometry was modified to include the HEFJR sting and
support in the hot and cold jet comparisons.
Figure 15: Capsule Pressure and Slice Locations, 40 Deg. nozzle
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Figure 16: Hot Jet CP, Cut 4, with and without HEFJR sting and capsule support, SST model
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Figure 17: Hot Jet CP, Cut 4, with and without HEFJR sting and capsule support, SST-CC model
III.B.5.
Turbulence Model Plume Comparison
A primary goal of this work is to validate the current Overflow simulation guidelines which recommend the
use of the Mentor SST turbulence model with no corrections for hot plume modeling. The other primary
goal is to validate the accuracy of Overflow simulations on hot flow plumes versus cold flow plumes. In
prior work, the standard (uncorrected) SST model gave better accuracy than other models in simulations of
experimental LAV-like flows. However, these prior experiments had only cold jet plumes, and most ongoing
and planned experiments also use cold jets.6
Fig. 18 gives work done by N. Gross4 that shows a comparison of the velocity magnitudes from experiment
and every turbulence model in Overflow, at several streamwise stations. The turbulence models available in
Overflow are:
• Baldwin-Barth (BB)
• Spalart-Almaras (SA)
• kω
• Menter SST model (SST)
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Figure 18: Comparisons of Overflow Turbulence Models to 85AA velocity magnitude.4
The results demonstrate that the SST model gives the best agreement with the PIV data, but even the
SST model gives mediocre accuracy. There are also several options to overflow’s SST model that affect its
performance when specific physical phenomena are present. The options of interest are the compressibility
correction (CC), high temperature correction (TC), and rotation and curvature model (RC), shown below
along with the Overflow 2.1 input flags.
• Menter SST model with no corrections [NQT=205, ICC=0, ITC=0, IRC=0]
• SST with Sarkar compressibility correction (SST-CC) [NQT=205, ICC=1, ITC=0, IRC=0]
• SST with Sarkar compressibility and Abdol-Hamid temperature corrections (SST-CTC) [NQT=205,
ICC=1, ITC=1, IRC=0].
• SST with Sarkar compressibility correction and Spalart-Allmaras rotation and curvature model (SSTCRC) [NQT=205, ICC=1, ITC=0, IRC=1].
An ad hoc modification of the CC modeling was also tested, in which the turbulence variable source terms
that produce the CC effect were scaled by 1/2, and the results are labeled “CC/2”. CFD solutions computed
with and without the CC modeling display too little and too much diffusion of the plume, respectively. This
behavior suggests that a weaker CC model might be appropriate. The CC/2 test is intended to demonstrate
the sensitivity of the results to a modification of the CC model. It is not proposed as a remedy to the
challenge of modeling turbulence in these flows.
The first set of simulations were run with Overflow version 2.0aa las, and the no-capsule 40 degree nozzle
configuration with the hot jet. The PIV cross-sectional slice locations for the 40 degree nozzle are shown in
Fig. 19, followed by axial slices of the PIV and CFD velocity magnitude contours in Fig. 20 and cross sectional
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slices of the PIV and CFD velocity magnitude and velocity vector components in Figs. 21 through 24. The
axial slice contours in Figs. 19 and 20 show the experimental PIV data as color contours, while the CFD
data in Fig. 20 is represented by black contour overlays. Comparing the axial slices for SST and SST-CC,
the SST-CC model produces the correct number of shock cells, while the SST model shock cell structure
decays too rapidly. The velocity magnitude contours in Fig. 21 also indicate the plume predicted with the
SST model decays too rapidly, but also illustrates the plume predicted with the SST-CC model has a shear
layer that is too thin with too much vorticity, leading to a skewed shape downstream. Also of note is the
unsteadiness of the plume downstream for the SST-CC model.
The U velocity, given in Fig. 22, is a sensitive measure of modeling accuracy because it reveals details
of the flow’s structure that cannot be seen in the velocity magnitude. For example, streamwise vorticity
that results from the interaction of the plume and crossflow appears as adjacent regions of opposite-signed
U velocity. In these U plots, red denotes flow to the right and blue to the left. The regions of positive and
negative U velocity on the tops of the plot are the freestream being diverted around the plume. Just below
there, are regions of inflow (toward the centerplane) that are the lower halves of the streamwise vortices
emerging from the plume/freestream interaction. This vortex “signature” is complicated by the shock-cell
structure. However, slices 4 through 6 show a dominant structure of outflow on top, with inflow below.
The U component of velocity shows some left-right asymmetry in the PIV data, which was also seen in the
velocity magnitude. Fine-scale features in the U field are very different in PIV, SST, and SST-CC plots
in most slices. At slice 1, there are small difference between SST, SST-CC, and PIV data, aside from the
asymmetry in the PIV data. At slice 2 through 6, there are significant differences between PIV and CFD
data in the structure and size of features in the flow. Two general observation are that (1) features in the
ICC = 0 solution are very “smoothed out” while the ICC = 1 results retain some of the non-smooth features
seen in the PIV data, and (2) the global size and velocity magnitude of the different U regions are better
predicted by ICC = 0
The V and W velocities given in Figs. 23 and 24, , mirror the trends of the velocity magnitude contours
with the ICC = 0 and ICC = 1 solution displaying too much and too little diffusion of the plume, respectively. In summary, Figs. 21 through 24 show the general shape of the experimental plume lies somewhere in
between the SST and SST-CC modeled plumes, and there are plume asymmetries and details of the velocity
fields that are not predicted well by either model.
Figure 19: 40 degree nozzle PIV cross-sectional slice locations
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Figure 20: 40 degree nozzle PIV axial slices, ICC=0 (top) and ICC=1 (bottom)
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Figure 21: Total Velocity, 40 deg. nozzle, hot jet – Experimental PIV, SST and SST-CC
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Figure 22: U Component of Velocity, 40 deg. nozzle, hot jet – Experimental PIV, SST and SST-CC
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Figure 23: V Component of Velocity, 40 deg. nozzle, hot jet – Experimental PIV, SST and SST-CC
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Figure 24: W Component of Velocity, 40 deg. nozzle, hot jet – Experimental PIV, SST and SST-CC
The next set of simulations were run with Overflow version 2.1ae, and the no-capsule 25 degree nozzle configuration with the hot jet. The SST, SST-CC, SST-CTC, and SST-CRC models, along with the ad hoc
CC/2 test were run for this configuration. The PIV cross-sectional slice locations for the 25 degree nozzle are
shown in Fig. 25, followed by axial slices of the PIV and CFD velocity magnitude contours in Fig. 26, and
PIV and CFD velocity magnitude and velocity component contour cross sections, including all the mentioned
models, in Figs. 27 through 30.
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As with the 40 degree nozzle, the velocity magnitude contours (Figs. 26 and 27) show the SST model
plume decays too rapidly while the SST-CC plume decays too slowly. The SST-CTC and SST-CRC models
show only small differences from the SST-CC model, but the SST-CC/2 test best predicts the experimental
plume. The U component for the 25 degree nozzle configuration (Fig. 28) shows more of the jet core
turbulence noted in the experimental data for the 25 degree nozzle case, with none of the CFD models
capturing the asymmetries. The V (Fig. 29) and W (Fig. 30) component miniwalls show less asymmetry
for the 25 degree nozzle experimental data than was seen in the 40 degree nozzle data, and the SST-CC/2
model remains the best simulation of the experimental plume.
In conclusion, the 40 degree nozzle hot jet case indicated neither the SST nor the SST-CC models correctly
predicted the plume shape, with the experimental plume shape roughly in between the two models. The 25
degree nozzle hot jet case confirmed this assumption with the SST-CC/2 model most closely predicting the
experimental plume shape and velocity components. However the U component of velocity for both cases
showed some asymmetries in the experimental plume that were not being predicted by any of the models.
Figure 25: 25 degree nozzle PIV cross-sectional slice locations
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Figure 26: 25 degree nozzle PIV axial slices, ICC=0 and ICC=1
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Figure 27: Total Velocity, 25 deg. nozzle, hot jet – Experimental PIV, SST, SST-CC, SST-CTC, SST-CC/2,
SST-CRC
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Figure 28: U Component of Velocity, 25 deg. nozzle, hot jet – Experimental PIV, SST, SST-CC, SST-CTC,
SST-CC/2, SST-CRC
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Figure 29: V Component of Velocity, 25 deg. nozzle, hot jet – Experimental PIV, SST, SST-CC, SST-CTC,
SST-CC/2, SST-CRC
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Figure 30: W Component of Velocity, 25 deg. nozzle, hot jet – Experimental PIV, SST, SST-CC, SST-CTC,
SST-CC/2, SST-CRC
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III.B.6.
Hot Jet Plume Modeling
Here we will compare the CFD cold jet and hot jet predictions with experiment to draw conclusions about
the validity of extrapolating to hot jet flows with CFD results.
Figures 31 through 34 show the capsule surface pressure contours and plume Mach contours for the SST
and SST-CC models for the 25◦ configuration. The surface pressure data and plume Mach contours shown
are plotted on a decimated mesh. Again the most noticeable differences between the models can be seen on
the top of the capsule. The SST-CC model shows a higher pressure at the end of the conic section and a
lower pressure at the nose and trailing edge of the capsule.
The SPIV setup was incapable of gathering useful data with the cold flow jet due to condensation, so we
will focus on the capsule pressure taps. The capsule pressure tap distribution was shown above in Figs. 3
and 4. We will be looking specifically at the circumferential tap rings corresponding to the cuts shown
in Fig. 15. The experimental results are averages of all runs completed for each configuration with error
bars constructed using equation (1). Figure 35 shows a sample of the experimental runs and the resulting
average and error bars, using the 25 degree nozzle case and the circumferential pressure distribution at cut
2. Included in Figure 35 are two CFD results for comparison.
Figures 36 through 43 show the circumferential pressure distributions corresponding to cuts 1-4 of the
hot and cold jet cases, side by side, for the 40 degree nozzle. Each plot includes curves for all possible
combinations of the SST turbulence model and corrections except for the SST-CC/2 model, which was only
run for the 25 degree nozzle. Each simulation includes the HEFJR sting and capsule support. None of the
curves match the experimental data at all locations, with the biggest discrepancies being at the top of the
capsule for cuts 1, and 2. There is a noticeably smaller gap between the CFD results and the experiment for
the cold jet simulation. The largest differences between CFD models for both the hot and cold jet simulations
correspond to the addition of the compressibility correction, although there is no CFD turbulence model
choice which best matches the experimental data for all slices.
Figures 44 through 51 show circumferential pressure distributions for the 25 degree nozzle with hot and
cold jet cases side by side. Again all plots includes curves for all possible combinations of the SST turbulence
model and corrections, and each simulation includes the HEFJR sting and capsule support. Given the
plume interacts more with the capsule at the 25◦ configuration, there is a more noticable difference between
turbulence models for this case. For cut 1, there again is a large discrepency between CFD at the top of
the capsule, which is greater for the hot plume case than the cold plume case. At cut 2, the circumferential
trend at the top side of the capsule changes between SST and SST-CC models with the SST model more
closely matching the hot plume trend, and SST-CC matching the cold plume trend. Cuts 3 and 4 show
similar trends between turbulence models, but with varying magnitudes at the top of the capsule. Again the
gap between CFD and experiment is smaller for the cold jet simulation than with the hot jet, but there is
no clear CFD turbulence model choice that consistently predicts the experiment.
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Figure 31: CFD ICC=0, 25 deg. Hot Jet, Plume Mach Contours, Cp Capsule Contours
Figure 32: CFD ICC=0, 25 deg. Hot Jet, Plume Mach Contours, Cp Capsule Contours
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Figure 33: CFD ICC=1, 25 deg. Hot Jet, Plume Mach Contours, Cp Capsule Contours
Figure 34: CFD ICC=1, 25 deg. Hot Jet, Plume Mach Contours, Cp Capsule Contours
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Figure 35: Experimental Averaging and Errors, 25 deg. nozzle Hot Jet CP, Cut 1
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Figure 36: 40 deg. nozzle Hot Jet CP, Cut 1
Figure 37: 40 deg. nozzle Cold Jet CP, Cut 1
Figure 38: 40 deg. nozzle Hot Jet CP, Cut 2
Figure 39: 40 deg. nozzle Cold Jet CP, Cut 2
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Figure 40: 40 deg. nozzle Hot Jet CP, Cut 3
Figure 41: 40 deg. nozzle Cold Jet CP, Cut 3
Figure 42: 40 deg. nozzle Hot Jet CP, Cut 4
Figure 43: 40 deg. nozzle Cold Jet CP, Cut 4
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Figure 44: 25 deg. nozzle Hot Jet CP, Cut 1
Figure 45: 25 deg. nozzle Cold Jet CP, Cut 1
Figure 46: 25 deg. nozzle Hot Jet CP, Cut 2
Figure 47: 25 deg. nozzle Cold Jet CP, Cut 2
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Figure 48: 25 deg. nozzle Hot Jet CP, Cut 3
Figure 49: 25 deg. nozzle Cold Jet CP, Cut 3
Figure 50: 25 deg. nozzle Hot Jet CP, Cut 4
Figure 51: 25 deg. nozzle Cold Jet CP, Cut 4
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IV.
Conclusions
At the conclusion of this work we have noted that of the four turbulence models available for use with
Overflow, the SST models performed the best in predicting the plume shape shown by the experimental
data. Though the Overflow SST models showed the most accurate plume predictions, neither the SST or
SST-CC model exactly predicts the plume shape shown by the experimental data. Plume shape predictions
using a compressibility correction scaled by one half show better results, but the good comparisons do not
translate to better predictions of capsule pressure. The lack of modeling all of the struts and supports did
impact the predictive capability of the CFD on the underside of the capsule, but not on the top where the
plume impinges on the capsule surface. Additionally, the PIV data of the U component of velocity showed
asymmetries pointing to more turbulence in the core of the jet than what the CFD is predicting.
The CFD predictions for the cold jet case did fall closer to the experimental data than CFD predictions
for the hot jet case for the 25 degree nozzle case. For the 40 degree case, however, the error bars on the
experimental data prevent a clear conclusion of the hot jet case being more difficult to model than the cold
jet. The largest discrepencies between experiment and CFD occur near the nose of the capsule on the top
side. The CFD overpredicts the pressure on the top of the capsule for the 40 degree case, and underpredicts
for the 25 degree case. We can conclude that an error in the extrapolation from cold jet plumes to hot jet
plumes with CFD does exist, and the results presented here indicate the magnitude of the errors vary with
angle of attack.
Acknowledgments
The work documented herein was performed in support of the NASA Orion Multi-Purpose Crew Vehicle
(MPCV). The authors would like to acknowledge Paul Stremel of STC and Ted Manning at NASA Ames
Research Center for their work generating the CFD grids and grid scripts used in this study.
References
1 Wernet, M., Wolter, J. D., Locke, R., Wroblewski, A., Childs, R., Nelson, A., “PIV Measurements of the CEV Hot Abort
Motor Plume for CFD Validation,” AIAA Paper, 48th Aerospace Sciences Meeting, Orlando, FL, 2010.
2 Menter, F.R., and Rumsey, C.L., Assessment of Two-Equation Turbulence Models for Transonic Flows, AIAA–94–2343,
June 1994.
3 Buning, P. G., Jespersen, D. C., Pulliam, T. H., Klopfer, G. H., Chan, W. M., Slotnick, J. P., Krist, S. E., and Renze, K.
J., Overflow Users Manual NASA Langley Research Center, Hampton, VA, 2002.
4 Gross, N., “Summary of 85-AA Experiment Computational Analysis,” Technical Report, NASA Johnson Space Center,
2009.
5 Bodonya, Daniel J., and Lele, Sanjiva K., “Current Status of Jet Noise Predictions Using Large-Eddy Simulation,” AIAA
J. Vol. 46, No. 2, 2008, p. 364.
6 Ross, J. C., and Brauckmann, G. J., Aerodynamic and Aeroacoustic Wind Tunnel Testing of the Orion Spacecraft, 29th
AIAA Applied Aerodynamics Conference, Honolulu, HI, 27-30 June 2011, American Institute of Aeronautics and Astronautics,
Reston, VA (submitted for publication)
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