Partially-Dressed Cavity-Closed Nose Landing Gear (PDCC-NLG)
Mehdi R. Khorrami
NASA Langley Research Center
Mail Stop 128
Hampton, VA 23681-2199
Tel: 757-864-3630
Email : mehdi.r.khorrami@nasa.gov
And
Thomas Van de Ven
Gulfstream Aerospace Corporation
Acoustics & Vibration Group
P.O. Box 2206, M/S R-06
Savannah, GA 31402-2206
Tel: 912-965-7509
Email: thomas.vandeven@gulfstream.com
1.0 Introduction
The following is a problem statement for the Category 4 benchmark problem in airframe
noise using a 1/4-scale Partially-Dressed Nose Landing Gear. This document describes
the context of the problem, experimental data available to validate both aerodynamic
flow and acoustic results, and criteria for meeting benchmark goals.
1.1 Problem Significance
A major goal of the aeroacoustics community is to mitigate the environmental impact of
noise associated with civil transports during take-off and landing operations. As a result
of continued progress in engine noise research, a significant portion of aircraft approach
noise is now generated by the airframe. Aircraft undercarriages are major contributors to
airframe noise. For landing gears, the close proximity of many bluff bodies of diverse
geometrical scales and shapes generates a highly complex and interactive unsteady flow
field. The flow field (including noise sources, noise generation mechanisms, etc.) and the
resulting acoustic field are poorly understood and predicted. Advances in prediction
methodologies for airframe noise require high-fidelity numerical simulations in
conjunction with high-quality experimental data to guide the development and validation
of physics-based mathematical models.
An extensive aeroacoustic database was acquired using a physical model of a Gulfstream
G550 aircraft nose landing gear. This database is to be used as a benchmark to validate
the predictive capabilities of current computational aeroacoustic (CAA) methodologies,
as well as guide future development of more capable CAA tools.
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1.2 Experimental Configuration
The tested model is a 1/4-scale high-fidelity replica of a Gulfstream G550 nose landing
gear that includes part of the lower fuselage section. To duplicate the angle-of-attack
effect during flight, inherent in the model design is an inclination angle of 3 degrees
relative to the streamwise direction (x axis). The nose gear geometry to be used for
benchmarking the computations is a “partially-dressed” version of the fully dressed gear
whereby the hydraulic and electrical lines, lighting cluster, and the steering assembly are
removed and the gear cavity is closed (taped over). The aeroacoustic measurements for
the partially-dressed cavity-closed nose landing gear (PDCC-NLG) were obtained in
multiple phases using two distinct wind tunnel configurations. Extensive aerodynamic
measurements were gathered with the model installed in the closed-wall Basic
Aerodynamic Research Tunnel (BART) at the NASA Langley Research Center (LaRC).
The corresponding acoustic, and limited aerodynamic, measurements were acquired in
the open-jet University of Florida Aeroacoustic Flow Facility (UFAFF).
The aerodynamic measurements (performed in BART) are available for M = 0.12, 0.145,
and 0.166. The data comprise steady and unsteady surface pressures plus planar particle
image velocimetry (PIV). Concurrent acoustic and surface pressure measurements
(conducted in UFAFF) are available for M = 0.145, 0.166, and 0.189. In both tunnels, the
flow over the shock-strut was tripped to ensure a fully turbulent flow. The other primary
gear components, mostly fabricated from polycarbonate material, can also be assumed to
have experienced turbulent flow due to the inherent surface roughness.
2. Problem Statement
Computational aeroacoustic simulations of the PDCC-NLG are solicited. Although
prediction of the acoustic field is the ultimate goal, submissions targeting only the near
field unsteady flow are also welcomed and accepted. Given the differences in the model
setup and tunnel configuration between the closed-wall tunnel (BART) and the open-jet
facility (UFAFF), submissions should follow one of the two configuration tracks
described below. Concurrent simulation of both setups is highly desirable and
encouraged. Such a dual attempt would provide a unique opportunity to fully scrutinize
and understand facility dependency of the database.
2.1 Track A Configuration
The configuration of interest under this track is the installed model in BART (Ref. 1) that
includes the tunnel test section. In this configuration, the model is mounted in an inverted
position on the floor of the tunnel (Fig.1). The BART test section is 28” (0.7112 m) high,
40” (1.016 m) wide, and 10’ (3.048 m) long. For reference purposes, the model height
from the tunnel floor to the top (bottom in aircraft coordinates) of the wheels is
approximately 16.8” (0.427 m) and the wheel diameter is 5.5” (0.1397 m). The tunnelfloor wall (from the entrance to the end of the test section) and the fuselage surface
should be treated as viscous boundaries. The tunnel side and ceiling walls can be
assumed to be inviscid surfaces with no detrimental effects on the fidelity of the
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simulations. However, the decision on whether to treat the three remaining tunnel walls
as inviscid or viscous surfaces is left to the individual parties.
z
y
x
Figure 1. Partially dressed nose landing gear model as tested in BART
Overall, simulation of the Track A setup is most appropriate for benchmarking the nearfield aerodynamic results as well as predicting the acoustic field using the time-resolved
pressure data on PDCC-NLG solid surfaces only. Acoustic predictions based on FW-H
integration that employ data from off-body permeable surfaces become questionable due
to reflection of the low- and mid-frequency acoustic waves from the tunnel walls.
2.2 Track B Configuration
This track is concerned with acoustic and surface pressure results based on the installed
model in the UFAFF (Ref. 2). Simulation of the full environment for the open-jet facility
that would include the incoming jet and the associated shear layers, collector, and the
surrounding anechoic chamber is beyond the scope of the current solicitation. As a first
step towards capturing the open-jet test environment, under Track B, simulation of the
PDCC-NLG in free field (bounded by a mounting wall on one side) is solicited. In this
configuration, the model is mounted from a ceiling wall which theoretically extends to
infinity along the transverse and aft directions (Fig. 2). In the experiment, the leading
edge of the fuselage (coincident with the coordinate system origin) is located 13.25
inches (0.3366 m) from the jet exit plane upstream of the model. However, it is left to the
individual parties to decide how much of the ceiling is captured within the computational
domain. Except for the ceiling, the computational domain contains no other tunnel walls.
Again, the ceiling wall plus the fuselage surface should be treated as viscous surfaces.
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y
x
z
Figure 2. Partially dressed nose landing gear model as tested in UFAFF
Simulation of the Track B setup is the more appropriate configuration for benchmarking
the intermediate distance acoustic field results. Validation of the acoustic field is to be
accomplished via the a) fluctuating pressure field on the gear surfaces, b) integration of
the FW-H equation using collected data at off-body permeable surfaces, or c) direct
simulation of the propagating acoustic waves. Attempts at acoustic comparisons utilizing
any combination of the three approaches are highly desirable and encouraged. For the
track B configuration, the corresponding validation of the simulated aerodynamic results
is accomplished via direct comparison with the steady and unsteady surface pressure data
available from the UFAFF tests. However, submitters are strongly encouraged to perform
cross validation checks with the more extensive aerodynamic data (i.e. planar PIV
dataset) available from the BART facility. Such comparisons would provide a unique
opportunity to fully scrutinize and understand facility dependency of the aerodynamic
database.
2.3 Track B (version 2) Configuration
Based on the insights gained from the results presented during the BANC-I workshop and
internal research at NASA, it has become clear that the Track A (BART) configuration is
quite unsuitable for far field acoustic prediction. Due to severe acoustic reflections from
the tunnel walls, the fluctuating pressure field on the solid surfaces, as well as the data on
FW-H permeable surfaces, produce results that are deemed ambiguous/unreliable at best.
A computational setup that has been developed and utilized internally by NASA and
found to produce relatively accurate far field acoustic results is a slightly modified
version of the UFAFF setup. The configuration, which is to be called Track B (version2),
consists of the Track B geometry with the flat plate acting as the ceiling, but now
arranged so that the entire model is suspended in free space. That is, as shown in Fig.3,
the free field computational domain is constructed such that air can flow both above and
below the flat plate (ceiling) on which the fuselage and nose landing gear are mounted. In
this configuration, the leading, side, and trailing edges of the flat plate are beveled to
eliminate or reduce local flow unsteadiness. The top surface of the plate may be
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considered as inviscid while the under surface of the plate accommodating the fuselage
and the nose gear must be treated as a viscous boundary, in line with the original Track B
setup.
a) side view
b) frontal view
Fig.3 Modified version of Track B configuration tailored for far field acoustic prediction
showing suspension of the flat plat and model inside the computational domain.
2.4 Provided Geometry Data
For the Track A setup, surface definition for the PDCC-NLG, including the BART test
section coordinates, will be provided in both STEP and IGES file formats. In this
configuration, the right-handed coordinate system is fixed at the leading edge of the
fuselage on the wind tunnel floor (Fig. 1) such that the x-axis corresponds to the
streamwise direction, the y-axis to the spanwise direction, and the z-axis to the vertical
(along the gear main strut) direction. In addition to the geometry file, the x-y-z
coordinates for the steady and unsteady surface pressure ports, and the locations of the
PIV planes selected for comparison with the computational results, will be provided.
For the Track B setup, surface definition for the PDCC-NLG, including the ceiling wall,
will be provided in both STEP and IGES file formats. To remain consistent with the
right-handed coordinate system of the closed-wall tunnel (Track A) configuration, the y
and z axes are rotated 180 degrees about the x axis (Fig. 2). Under this transformation,
the coordinates for the surface pressure ports and the locations of the PIV planes remain
identical between the Track B and Track A configurations.
For the Track B (version 2) setup, surface definition for the PDCC-NLG, including the
beveled flap plate geometry acting as the ceiling wall, will be provided in both STEP and
IGES file formats. All other aspects of this revised setup remain the same as in the above
paragraph for the Track B configuration.
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The linear array microphones of interest (M2 through M10) are roughly 45” (1.143 m) to
56” (1.422 m) from the wheels’ axle (Fig. 4). The linear array shear layer corrections
were performed based on the assumption that the “noise source” is effectively localized
at the location of the upper torque arm pressure sensor (Fig.4). The sensitivity of the
corrections to the source location will be investigated and reported in the future. For the
purpose of acoustic comparisons, the coordinates for the individual microphones of the
linear array in both flyover and sideline orientations will be furnished. For each
orientation, two sets of coordinates corresponding to the physical and virtual (corrected
position due to the presence of the open-jet shear layers, as shown in Fig. 4) locations of
the microphones will be provided.
=Re
M6'
M6
Figure 4. Linear array flyover configuration in UFAFF with shear
layer correction nomenclature.
Shear layer correction (SLC) nomenclature:
Rm:
Measured source-to-microphone distance
θm:
Measured source-to-microphone angle
Re:
Effective source-to-microphone distance
θc :
Corrected propagation angle
h:
Source-to-shear layer distance
Z:
Distance from source to plane of microphones
M6′: the corresponding virtual position for microphone 6 in the flow field. M6′
preserves the directivity of the source and the ray path length for the virtual position of
the microphone in the velocity field.
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2.5 Benchmark Flow Conditions
Given that the free stream Mach numbers of 0.145 and 0.166 were common to the two
tunnels, the dataset at the M=0.166 speed regime is selected as the benchmark case for
aeroacoustic simulations of both Track A and B configurations.
2.5.1 BART Conditions
The flow Reynolds number based on the diameter of the shock strut (0.75” or 0.01905 m)
and the free stream Mach number of 0.166 (185.7 ft/s or 56.6 m/s) is 73,000. The total
pressure and temperature in the BART settling chamber (or the inflow plane) are 14.716
psi (101,464 Pa) and 73.9 degrees F (23.28 C) respectively, and the measured static
pressure at the exit plane is 14.3936 psi (99,241 Pa). The free stream turbulence intensity
(TI) at this Mach number is 0.077%. The boundary layer on the tunnel floor at the
entrance to the test section possesses a shape factor that is between 1.3 and 1.4 and thus
should be assumed as fully turbulent. On the centerline at the test section entrance, the
boundary layer and the displacement thicknesses are 0.61 inches (1.55 cm) and 0.086
inches (0.22 cm), respectively.
2.5.2 UFAFF Conditions
Due to the higher temperature and humidity of the air, the free stream speed in UFAFF
corresponding to M=0.166 is slightly higher (188.3 ft/s or 57.4 m/s) than that of BART.
Using a viscosity of = 1.895e-5 kg/m sec, the flow Reynolds number based on the
diameter of the shock strut (0.75” or 0.01905 m) and the free stream Mach number of
0.166 is 66,000. The free stream static and dynamic pressures at the jet exit (or the
inflow) plane are 14.3 psi (98,605 Pa) and 0.2756 psi (1,900 Pa), respectively. The tunnel
ambient pressure is 14.59 psi (100,595 Pa) and the temperature is 93 degrees F (33.9 C).
The density of the humid air is 1.14 kg/m3. The free stream turbulence intensity (TI) at
this Mach number is < 0.1%. The boundary layer thickness on the wall at the jet exit
plane is approximately 0.9 inches (2.29 cm).
3. Simulation Roadmap
Comparison of simulation results with the available aeroacoustic measurements (or even
with results from other codes) is typically quite challenging. In the absence of certain
rules and roadmaps on how to process and report the CAA results, such comparisons
have the potential of becoming the source of extra confusion rather than an insightful
exercise. It is not the intention of the current solicitation to dictate the specifics of any
technical approach. However, the roadmap described below attempts to 1) provide
minimal constraints, and 2) achieve uniform reporting guidelines for all submitting
parties.
For the remaining sections of this document, it is important to emphasize that all
participating authors must adhere to the requirements highlighted in italic font when
reporting their results. In addition, certain parameters/results, highlighted in blue font,
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must be provided to the PDCC-NLG POC (Mehdi Khorrami) by Thursday May 29, 2014
prior to the Workshop. Authors who fail to report/furnish required metrics will be
prevented from presenting their results at the workshop. The collected data will be
used in an anonymous fashion to discuss code to code comparisons and the overall trends
of the submitted computational results.
3.1 Time Resolution and Unsteady Data Collection
The sampling frequency for surface pressure measurements is 51.2 kHz yielding a
bandwidth of 20 kHz (the A/D converter employed has a Nyquist factor of 2.56). The
measured data records are 32 seconds long containing 1,638,400 samples. The data
records are broken into 100 blocks of 16384 samples to construct the power spectral
density (PSD) at each sensor location. The resulting frequency bin width for PSD is 3.125
Hz (=100/32s). The Prms is obtained from integration of the PSD curves over the 3.12520,000 Hz bandwidth. However, given the experience gained from the BANC-I results
and the extreme challenges of generating computational time records that are of equal
duration to the measurements, comparisons of measured vs. simulated PSDs are restricted
to the 400 Hz – 10 kHz frequency range. Authors wishing to extend such comparisons to
lower frequencies are encouraged to do so, but this is optional.
While resolution of frequencies up to 16 kHz is of great interest, the submitting parties
should maintain at a minimum a frequency resolution of 8-10 kHz for the unsteady flow
field. In addition, the simulated data records should be of sufficient duration to allow the
construction of the surface pressure PSD curves based on averaging of three independent
data blocks using a frequency bin width that is equal to or smaller than 16 Hz.
Accordingly, the post transient simulated data records must be 0.19 seconds or longer to
allow a minimum of three data blocks of 0.0625 seconds to be constructed. However,
generation of longer data records is highly encouraged. The choice of windowing and
data overlap is left to the authors.
3.2 Transient Data
For highly interactive and complex flows such as those encountered over landing gears,
the transient flow is of significant duration and can easily be mistaken with the expected
final well-established unsteady flow. Inclusion of this initial transient segment in both the
sampled time records and the construction of the mean quantities irrevocably corrupt the
computed results.
Submitting authors must ensure that upon the start of time-accurate simulations, at a
minimum, the first 0.06 seconds of the computed data are discarded and not used for any
post processing purposes. Based on free stream M=0.166, the 0.06s period is equivalent
to roughly 1.4 times the convective time-scale for an eddy to travel from the wheel axle
location to the exit plane of the BART test section. Alternatively, the 0.06s period can be
viewed as equivalent to 24.5 times the convective time-scale for eddies to traverse one
wheel diameter of 5.5” (0.1397 m). However, authors should ensure that sufficient time is
allowed for their CAA tools to fully converge prior to acquiring data for post processing.
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3.3 Convergence Metrics
The convergence behavior of the simulated flow field is of paramount importance and
should be clearly demonstrated by every submitting author. As a global indicator, average
lift and drag forces for the nose gear, without the contribution from the fuselage, must be
tabulated and reported. Here, the term lift corresponds to the force (FZ) along the Zcoordinate and the term drag (FX) signifies the force along the X-coordinate. Use an area
of 0.06945 ft2 (0.006452 m2) to obtain the lift (FZ) and drag (FX) coefficients. (Additional
reporting of lift and drag on some gear subcomponents such as wheels, door, torque arm,
etc. is also encouraged as relevant metrics but is considered optional at this time). In
addition, documentation and reporting of the averaging time required to obtain timeaveraged and fluctuating pressures is required. To clearly demonstrate convergence, the
authors should extend the time duration of the initial computational record by 40%,
reprocess the data, and then show the percentage of change in the selected metrics.
Demonstration of grid convergence is highly desirable but difficult to achieve. In
addition to the finest resolution possible, all authors must run their simulations at one or
more (coarser) resolutions and report the incurred changes. If feasible, the Grid
Convergence Index (GCI) approach explained in chapter 5, page 114 of Ref. 3 could be
utilized to highlight the grid sensitivity of the computed results.
3.4 Other Metrics
All submissions must provide the total grid count (e.g. number of nodes, cells, etc.) and
the cell size nearest to the gear surface at select locations on the shock strut, starboard
wheel, and torque link. For track B participants, the dimensions of the computational
domain utilized must also be reported. Additional data regarding the computational
resources that need to be furnished by the authors are: a) the type and number of CPUs
utilized, b) the type of data communications used, c) the time required to obtain 1000
time steps, and d) other relevant parameters unique to the computational resources
utilized.
4. Comparison with Measurements
Due to the extensive amount of data collected, benchmarking and validation of the
simulations will be limited to a select subset of the aeroacoustic database. The type and
locations of selected data are highlighted in the following subsections. Submitting authors
should provide a full account of the comparison between their simulations and the
highlighted data.
4.1 Time-Averaged Surface Pressures (Cp)
One circumferential (0 to 360 degrees) row on the starboard wheel, two transverse rows
on the port wheel, and five spanwise rows on the door.
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4.2 Fluctuating Surface Pressures
Cprms and power spectral density (PSD) for all nine sensors (two on the starboard wheel,
one on the torque links, three on the door, two on the drag link, and one on the main
strut). Additional PSD data from the roving (mobile) sensor, installed at various gear
locations, can be made available for qualitative comparisons only. Sample fluctuating
pressure (p′) time history (1000-1500 time steps) at four selected probe locations (one on
starboard wheel (sensor 7), one on torque-link (sensor 15), and two on door (sensors 4
and 10)) should be provided. The Cprms value for all nine sensors, obtained from
integration of the PSD plot between 400 Hz and 10 kHz, must be provided to the POC for
category 4 submissions by May 29, 2014.
4.3 Planar PIV Data
The planar PIV data for validating simulation results are comprised of three x-y planes
adjacent to the starboard wheel, three x-y and three x-z planes in the wake of the wheels,
one x-y plane in the wake of the torque-link, and three x-y planes in the wake of the door.
Inherent in the PIV data are the two in-plane mean velocity components, the first and
second moments of the fluctuating velocity components, and the 2D turbulent kinetic
energy (TKE). At a minimum, authors should provide comparisons of the contours for
the two mean velocity components and the 2D TKE.
The authors must provide a profile of the averaged 2D TKE, Z-vorticity component, and
streamwise velocity, along four selected line cuts. The coordinates for the line cuts are
included in the experimental database to be made available to all interested parties.
4.4 LDV Data
LDV measurements of the regions that are hard to map using the PIV technique (i.e.,
between the two wheels, main strut and door, etc.) are to be obtained at UFAFF in 2011.
The data will be disseminated to the interested parties as soon as they become available.
4.5 Acoustics
Acoustic intensity and spectra based on linear array measurements (nine microphones) at
flyover and sideline orientations.
The authors may provide acoustic spectra at microphone locations M4, M5, M7, M8, and
M9, with M7 being the most important. For code to code comparison, the spectra must be
submitted as Power Spectral Density (PSD) in [psi2/Hz]. For comparison with
experimental measurements, 1/3rd Octave Sound Pressure Levels (SPL) in dB must be
provided at the same microphone locations.
5. Code to Code Comparison
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As stated throughout this document, the participating authors must submit the following
information by May 29, 2014 ahead of the BANC-II workshop:
a) Average lift and drag forces for the nose landing gear without the contribution
from the fuselage. The term lift corresponds to the force (FZ) along the Zcoordinate and the term drag (FX) signifies the force along the X-coordinate. The
lift (FZ) and drag (FX) coefficients should be computed using an area of 0.06945
ft2 (0.006452 m2).
b) Cprms values for all nine sensors. The reported Cprms data must be computed from
integration of the PSD curves between 400 Hz and 10 kHz
c) Profiles of averaged 2D TKE, Z-vorticity component, and streamwise velocity at
the four specified line cut locations.
d) Power Spectral Density (PSD) in [psi2/Hz] and 1/3rd Octave Sound Pressure
Levels (SPL) in dB at microphones M4, M7, and M9.
6. References
1. Neuhart, D.H., Khorrami, M.R., and Choudhari, M.M., “Aerodynamics of a
Gulfstream G550 Nose Landing Gear Model,” AIAA Paper 2009-3152, Miami, May 1113, 2009.
2. Zawodny, N.S., Liu, F., Yardibi, T., Cattafeta, L.N., Khorrami, M.R., Neuhart, D., and
Van de Ven, T., “A Comparative Aeroacoustic Study of a ¼-Scale Gulfstream G550
Aircraft Nose Landing Gear Model,” AIAA Paper 2009-3153, May 2009.
3. Roache, P. J., Verification and Validation in Computational Science and Engineering,
Published by Hermosa Publishers, 1998.
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