FLIGHTSTREAM - OpenVSP

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

http://www.researchinflight.com/

http://www.researchinflight.com/

https://researchinflight.com/contact us.html

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)


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



0.0

0.5

1.0

1.5

2.0

-4 0 4 8 12 16 20 24

CL


Data



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



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



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


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○


FLOW SEPARATION



Mach 0.2

AOA = 30○


FLOW SEPARATION



Mach 0.2

AOA = 30○

Inviscid Viscous




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



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


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



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

Documents

FLIGHTSTREAM - OpenVSP