flightstream - openvsp
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FLIGHTSTREAM®
VISCOUS MODELING & RESULTS
• Established 2012
• Primary functions are the development, marketing and support of FlightStream and the development of aerodynamic solutions
• Website: https://researchinflight.com
RESEARCH IN FLIGHT COMPANY
• Website: https://researchinflight.com
• Contact us page
FOUNDERS
• 670 North College Street, Suite A, Auburn, Alabama 36830
LOCATION
FLIGHTSTREAM ®
• FlightStream® is a highly efficient subsonic, inviscid, surface-vorticity flow solver
• Capabilities:
No volume meshes needed
High Fidelity Inviscid Load Calculations for Airplanes of a wide variety of configurations including blended bodies, canard configurations, and nearly any nonconventional geometry.
Generates accurate results in minutes.
Industry validated across a range of geometries and applications.
CAD-based geometry import and surface meshing.
Highly intuitive and scriptable UI & High Quality native graphical post processing.
Gas Turbine engine integration through NPSS, inlet definition, and exhaust modeling.
INTRODUCTION
CAD
High geometric fidelity.
High-quality anisotropic mesh.
CAD-based physics.
Higher one-time setup time.
COMPGEOM
Anisotropic mesh.
Simple one-click operation.
Fast.
User needs to return to VSP to make mesh changes.
CFD-MESH
Isotropic mesh only.
Good-quality mesh.
Integrated with VSP: good geometric fidelity of meshes.
Slower.
OPENVSP: ROUTES TO FLIGHTSTREAM ®
INTRODUCTION GEOMETRY
• Streamlines
• Surface streamlines
• Off-body streamlines
• Stream tubes
• Stream line distributions
• 3D modeling of streamlines
• Growing streamlines from probe points
• Upstream/Downstream growth
• Flow contours along streamlines
• Probe points
• User-specified probing locations in 3D-space
• Import/Export spreadsheet of probe point clouds
• Probe surfaces
• Generate a cloud of probe points from individual components
• Sectional planes
• Pressure and Mach number contours
ANALYSIS CAPABILITIES
INTRODUCTION
• FlightStream® is completely scriptable in a command line format.
• Users can execute a scripted FlightStream® run with a specified simulation settings and geometry files.
• FlightStream® scripting is text-file based.
SCRIPTING
INTRODUCTION
VISCOUS MODELS
• A de-coupled, generalized, compressible flow-separation model has been implemented into FlightStream.
• Separation line contours.• Laminar separation bubble capability.• Post-stall aerodynamic loads and moments.• CLMAX and stall angle predictions.
FLOW SEPARATION
• Automated on-body streamlines and BL transition models have been added as an essential feature in computing:
• Skin-friction models using laminar and turbulent profiles.• Post-processing capabilities along user-defined streamlines.• Transition line contours.
• Applications• Compute inviscid boundary distortion.• Advanced skin-friction models.• Propulsion integration.• Boundary Layer Ingestion Toolbox.
FLOW SEPARATION
Flow Separation
On-body Streamlines
Boundary Layers
Transition Model
Separation Marker
Post-separation
Velocity
Post-separation
forces
FLOW SEPARATION
BOUNDARY LAYER INPUTS
The following input parameters are required from the user for generating the
FlightStream integral boundary layers:
• Boundary layer type (Laminar or Turbulent)
• Velocity profile choices (Power-law or Coles)
• Power law exponents (or auto-compute)
• Coles scaling factor
• Initial shape factor
• Wall temperature
All inputs are specified in the fluid properties node (simulation tree) or in the
Analysis tab interface.
BOUNDARY LAYERS
BOUNDARY LAYER: NACA 0012
Mach 0.3 flow at 0○ AOA
0.0
0.2
0.4
0.6
0.8
1.0
1.2
1.4
0.0 0.2 0.4 0.6 0.8 1.0
Bo
un
da
ry L
ay
er
Thic
kn
ess
(in
.)
x/c
Boundary Layer Thickness (in.)
FlightStream (Turbulent compressible)
NACA Wind Tunnel
Becker, J., “Boundary Layer Transition of the NACA 0012 and 23012 Airfoils in the 8-
Foot High-Speed Tunnel”, NACA Wartime Report, January 1940
BOUNDARY LAYERS
VALIDATION STUDIES: MIT/NASA D8Surface Pressure contours on the MIT D8 in FlightStream®.
𝛼 = 2°,𝛽 = 0°, 𝑉∞ = 70 𝑚𝑝ℎ. Engine fan faces modeled as velocity inlets at 𝑉𝑖𝑛𝑙𝑒𝑡 = 30 𝑚𝑝ℎ. 𝑈𝑡𝑖𝑝 𝑉∞ = 2.64
Streamline measured from centerline of propulsor inlet
boundary and computed TCBL data from FlightStream
0.00
0.02
0.04
0.06
0.08
0.10
0.12
0.14
0.000
0.005
0.010
0.015
0.020
0.025
0.0 0.5 1.0 1.5 2.0 2.5
Ma
ch
Nu
mb
er
B. L.
Th
ick
ne
ss (
m)
X (Meters; Measured From Aircraft Nose)
FlightStream (Laminar)
FlightStream (Turbulent)
Mach Number
BOUNDARY LAYERS
INLETS & EXHAUST JETS• Surfaces can be marked as velocity or mass-flow inlets
• Can be coupled with NPSS to create integrated engine simulations
• Created using local coordinate systems
• Need only radius, exhaust velocity and fluid parameters as user inputs
• Can be cascaded to model concentric jets
BOUNDARY LAYERS
VALIDATION STUDIES: MIT/NASA D8
Surface Pressure contours on the MIT D8 in FlightStream®.
𝛼 = 2°,𝛽 = 0°, 𝑉∞ = 70 𝑚𝑝ℎ. Engine fan faces modeled as velocity inlets at 𝑉𝑖𝑛𝑙𝑒𝑡 = 30 𝑚𝑝ℎ. 𝑈𝑡𝑖𝑝 𝑉∞ = 2.64
Inlet Viscous Distortion Maps (Coefficient of total pressure
contours) for the starboard propulsor on the MIT D8.
NASA LaRC experimental measurements FlightStream®
Uranga, A., Mark Drela, et.al., “Preliminary Experimental Assessment of the Boundary Layer Ingestion Benefit for the D8 Aircraft,” AIAA 2014-0906
BOUNDARY LAYERS
RESULTS
• A variety of basic geometries and airplane configurations have been tested for this effort.
RESULTS
RESULTS: SPHERE
• Separation markers at Re=4.25 Million:
Forced laminar Turbulent
RESULTS
-1.5
-1.0
-0.5
0.0
0.5
1.0
1.5
-0.5 0.0 0.5
CP
X
FlightStream (Potential flow)
Experiment
FlightStream (Separation)
RESULTS: SPHERE
Shoemaker, J. M., “Preliminary Biplane Tests in the Variable Density Wind Tunnel,” NACA Technical Note No. 289, NACA, 1928
RESULTS
RESULTS: BIPLANE WINGS
Incompressible flow, 0° Stagger wings.
Shoemaker, J. M., “Preliminary Biplane Tests in the Variable Density Wind Tunnel,” NACA Technical Note No. 289, NACA, 1928
FLOW SEPARATION
-0.5
0.0
0.5
1.0
1.5
2.0
2.5
-10 10 30 50
CL
Angle of Attack (Deg)
Shoemaker Data, 1928
FlightStream (Linear)
FlightStream (Separation)
RESULTS: NASA EET-AR-12
FLOW SEPARATION
0.0
0.2
0.4
0.6
0.8
1.0
1.2
1.4
1.6
1.8
0.00 0.20 0.40 0.60
CL
CDData
FlightStream (Linear)
FlightStream (Separation)
0.0
0.5
1.0
1.5
2.0
-4 0 4 8 12 16 20 24
CL
Angle of Attack (Deg)
Data
FlightStream (Linear)
FlightStream (Separation)
Data: Olson E.D., and Albertson C.W., "Aircraft High-Lift Aerodynamic Analysis Using a Surface-Vorticity Solver", 54TH AIAA Aerospace Sciences Meeting, AIAA SciTech Forum, (AIAA 2016-0779)
RESULTS: NASA EET-AR-12
Data: Olson E.D., and Albertson C.W., "Aircraft High-Lift Aerodynamic Analysis Using a Surface-Vorticity Solver", 54TH AIAA Aerospace Sciences Meeting, AIAA SciTech Forum, (AIAA 2016-0779)
FLOW SEPARATION
-1.2
-1.0
-0.8
-0.6
-0.4
-0.2
0.0
0.2
-4 -2 0 2 4 6 8 101214161820222426
CM
Angle of attack (Deg)
AVL (ideal)
AVL + DATCOM
Data
FlightStream (Linear)
FlightStream (Separation)
0.0
0.2
0.4
0.6
0.8
1.0
1.2
1.4
1.6
1.8
-0.6 -0.4 -0.2 0.0 0.2
CL
CM
RESULTS: DLR-F4
FLOW SEPARATION
-0.1
0.1
0.3
0.5
0.7
0.9
1.1
1.3
1.5
0.00 0.05 0.10 0.15 0.20
CL
CD
NLR HST
ONERA-S2MA
DRA 8ft x 8ft
FlightStream (Linear)
FlightStream (Separation)
Mach 0.6
RESULTS: CESSNA-210 NLF MOD.
Murri, D. G. et.al, “Wind Tunnel Results of the Low-Speed NLF(1)-0414F Airfoil”, Research in Natural Laminar Flow Control, Part 3, p 673-696, December, 1987
FLOW SEPARATION
-0.5
0.0
0.5
1.0
1.5
2.0
2.5
-10 0 10 20 30
CL
Angle of Attack (Deg)
NASA Langley WT DataFlightStream (Linear)FlightStream (Separation)
RESULTS: NASA TRAP WING
FLOW SEPARATION
AIAA HLPW-1 Geometry
Reynolds Number = 4.3 Million
Mach 0.2
AOA = 30○
RESULTS: NASA TRAP WING
FLOW SEPARATION
AIAA HLPW-1 Geometry
Reynolds Number = 4.3 Million
Mach 0.2
AOA = 30○
RESULTS: NASA TRAP WING
FLOW SEPARATION
AIAA HLPW-1 Geometry
Reynolds Number = 4.3 Million
Mach 0.2
AOA = 30○
Inviscid Viscous
RESULTS: NASA TRAP WING
AIAA HLPW-1 Geometry
Reynolds Number = 4.3 Million
Mach 0.2
AOA = 30○
Experiment results from the 2ND AIAA CFD High-Lift Prediction Workshop, San Diego, California, June 2013
FLOW SEPARATION
0.50
1.00
1.50
2.00
2.50
3.00
3.50
4.00
-5 0 5 10 15 20 25 30 35 40 45
CL
Angle of Attack (Degrees)
HLPW1 Experiment
FlightStream (Separation)
FlightStream (Linear)
0.5
1.0
1.5
2.0
2.5
3.0
3.5
4.0
4.5
5.0
0 4 8 12 16 20
CL-CLSTAB
Angle of Attack (Deg)
FlightStream (viscous)
FUN3D
RESULTS: NASA X-57
FLOW SEPARATION
X-57 + HL Nacelles
X-57 + HL Nacelles + Flaps
X-57 + HL Nacelles + Power + Flaps
DRAG: NASA X-57
FLOW SEPARATION
0.0
0.5
1.0
1.5
2.0
2.5
3.0
3.5
4.0
0.00 0.05 0.10 0.15 0.20 0.25
CL
CD
FUN3D
FlightStream (Linear)
FlightStream (Separation)
CURRENT ACTIVITIES
• Research in Flight is working with Skyborne Technology in Port St. Joe, Florida on a Phase II Extended Option to enable and expand the Phase II deliverables to Airship Design.
CURRENT ACTIVITIES
CURRENT ACTIVITIES
PHASE II
Coupled integral boundary layer
Post separation base drag
Propeller actuator BLI effects
Propeller actuator effects to flow
separation
High incidence BLI
FlightStream licenses for NASA
Demonstration of technology on SMA 600 class
airship
User Documentation
Validation studies
(NASA + Skyborne Technology)
Phase II-E
CURRENT ACTIVITIES
ACKNOWLEDGEMENTS
• NASA Small Business Innovative Research (SBIR) contract award, Phase I (2016). “Robust prediction of high lift using surface vorticity”.
• NASA Small Business Innovative Research (SBIR) contract award, Phase II (2017). “Robust prediction of high lift using surface vorticity”.
• NASA Small Business Innovative Research (SBIR) contract award, Phase II E (2019). “Robust prediction of high lift using surface vorticity”.
• NASA Small Business Technology Transfer (STTR) contract award, Phase I (2019). “Air Vehicle Gust Response Analysis for Early Design”.
SUMMARY
THANK YOU
www.researchinflight.com