Activity sheet airfoil and wing fluent simulation

ACTIVITY SHEETAIRFOIL

FLUENT SIMULATIONAE4015 COMPUTATIONAL AERODYNAMICS

EMA AMALIA, ST., MT.

SIMULATION OF AIRFOIL NACA 65(2)-415

PREPARING GEOMETRY OF THE AIRFOIL

• TO DO SIMULATION ON AIRFOIL NACA 65(2)-415, HERE WE WILL USE ICEM CFD TO

CONSTRUCT GEOMETRY

• AIRFOIL THAT HAS BEEN CONSTRUCTED USING SPACECLAIM IS FOR CONSTRUCTING WING. IT

WON’T BE GOOD FOR AIRFOIL CFD SIMULATION

• NOW WE WILL CONTINUE, FIRST, WE DOWNLOAD THE AIRFOIL COORDINATE FROM “UIUC

AIRFOIL DATA SITE” UNDER THE ALPHABET “N” (HTTPS://M-

SELIG.AE.ILLINOIS.EDU/ADS/COORD_DATABASE.HTML#N) AND CHOOSE “NACA652415.DAT”

• OPEN THE FILE “NACA652415.DAT” IN MS EXCEL, ADD AN ADDITIONAL COLUMN FOR Z VALUE OF 0 (BECAUSE

WE WORK IN 2 DIMENSIONAL).

• ADD SOME INFO AT THE TOP OF FILE AS IN THE NEXT SLIDE

• REMEMBER THAT THE SEQUENCE OF AIRFOIL COORDINATE DATA SHOULD CONSIST OF TWO CURVES:

1. FROM LEADING EDGE TO TRAILING EDGE THROUGH UPPER SURFACE

2. FROM LEADING EDGE TO TRAILING EDGE THROUGH LOWER SURFACE

• REMEMBER TO ADJUST THE NUMBER OF POINTS TO MATCH WITH THE NUMBER OF COORDINATE

DATA POINTS. FOR EXAMPLE, IF YOU WRITE DOWN THE DATA NUMBER IS 26, YOU SHOULD MODIFY

THE COORDINATE DATA POINTS TO BE 26 TOO.

Original

NACA652415.DAT

Number of point

in each curve

Number of curves

Change sequence of upper surface

to be from LE to TE

Adding a point of 0,0,0 to upper

surface so that the number of point

will be 26 for each curves

Additional column of z

• COPY THE MODIFIED DATA OF AIRFOIL COORDINATE TO A NEW SHEET, AND SAVE IN FORMAT

“TEXT (TAB DELIMITED)” AS *.DAT”, FOR EXAMPLE “NACA652415PROFILE.DAT”. BECAUSE EXCEL

HAS NO FORMAT OF “*.DAT”, JUST ENTER “NACA652415PROFILE.DAT” AS INPUT TO FILE NAME

• NOW, THE *.DAT FILE IS READY TO BE IMPORTED FROM ICEM CFD

• LAUNCH ICEM CFD:

1. FROM WORKBENCH: DRAG “ICEM CFD” FROM LEFT WINDOW TO MAIN WINDOW AND

RIGHT CLICK “MODEL”, CHOOSE “EDIT”. THEN ICEM CFD WILL BE LAUNCHED. OR DOUBLE CLICK

“MODEL”, AND ICEM CFD WILL BE LAUNCHED.

• OR LAUNCH ICEM CFD STANDALONE

• THE DIFFERENCE BETWEEN ICEM CFD LAUNCHED FROM WORKBENCH AND STANDALONE

Launched from Workbench

Launched Standalone

• IF YOU WORK WITH WORKBENCH, DON’T FORGET TO “SAVE PROJECT” FROM FILE TAB (DO

NOT USE UPDATE PROJECT), EXIT TO WORKBENCH, AND SAVE YOUR WORKBENCH PROJECT IN

DIFFERENT NAMES FOR SOME STEPS, IN CASE YOU DO A WRONG STEP, YOU DON’T NEED TO

REPEAT FROM ZERO.

• IF YOU WORK WITH ICEM CFD STANDALONE, DON’T FORGET TO “SAVE PROJECT AS”

FREQUENTLY WITH DIFFERENT NAMES FOR SOME STEPS, SO THAT IF YOU DO A WRONG STEP,

YOU DON’T NEED TO START FROM ZERO.

(YOU CAN DO THIS FROM “FILE” TAB, AND CHOOSE “SAVE PROJECT AS”, THEN WRITE DOWN

THE NAME OF INTERMEDIATE FILE, FOR EXAMPLE: “NACA652415 STEP 1. PRJ”

• IN THIS ACTIVITY SHEET, THE GUIDELINES WILL BE MAINLY DONE FROM WORKBENCH. BUT IF

YOU WORK FROM STANDALONE ICEM CFD, IT WILL BE VERY SIMILAR.

• FROM ICEM CFD, CHOOSE FILE => IMPORT GEOMETRY => FORMATTED POINT DATA. SELECT

THE *.DAT FILE “ (FOR EXAMPLE: NACA652415PROFILE.DAT)

• THEN CLICK APPLY IN LOWER-LEFT WINDOW

• IF WE CHECK “POINTS” IN STRUCTURE TREE, THEN POINTS OF AIRFOIL COORDINATE WILL

APPEAR

PREPARING GEOMETRY OF THE AIRFOIL• RENAME THE PARTS. FOR EXAMPLE: CRVS BECOMES AIRFOIL BY RIGHT CLICK THE PARTS

• DELETE UNNECESSARY PARTS, FOR EXAMPLE “SURFS”

MAKING COMPUTATIONAL DOMAIN

• NOW WE WILL MAKE COMPUTATIONAL DOMAIN

• WE WILL CREATE SOME POINTS AS FOLLOWS:

(-20,0,0); (0,20,0); (0,-20,0); (21,20,0), (21,-20,0)

MAKING COMPUTATIONAL DOMAIN• FROM GEOMETRY TAB, CHOOSE “CREATE POINT”, CHOOSE “XYZ” OPTION TO INPUT COORDINATES

Uncheck

XYZ option

Enter coordinates

of points

MAKING COMPUTATIONAL DOMAIN• REPEAT TO DRAW POINTS OF OUTER BOUNDARIES OF COMPUTATIONAL DOMAIN

(-20,0,0)

(0,20,0)

(0,-20,0)

(21,20,0)

(21,-20,0)

MAKING COMPUTATIONAL DOMAIN

• CREATE CURVE FROM GEOMETRY TAB, CONNECT THE OUTER BOUNDARIES OF

COMPUTATIONAL DOMAIN

Uncheck

Choose this option

for straight liines

Click to select a point, middle click to confirm,

Repeat for all straight lines

MAKING COMPUTATIONAL DOMAIN• CREATE CURVE FROM GEOMETRY TAB, CONNECT THE OUTER BOUNDARIES OF

COMPUTATIONAL DOMAIN

Uncheck

Choose this option

for half circle

Click to select three points that construct semi-circle

Boundary, middle click to confirm

MAKING COMPUTATIONAL DOMAIN• WE WILL MAKE SURFACE OF COMPUTATIONAL DOMAIN

• CLICK “CREATE SURFACE” IN GEOMETRY TAB, AND CHOOSE “SIMPLE SURFACE” SELECT CURVES

OF OUTER BOUNDARIES BY CLICKS AND MIDDLE CLICKS TO CONFIRM

Enter the name of part as “Fluid”

Uncheck

Choose this simple surface option

CREATE BLOCKING• CHOOSE BLOCKING TAB, CLICK CREATE BLOCK

• CHOOSE PART OF “FLUID” AS BLOCKING, AND SELECT “2D PLANAR”, CLICK APPLY

CREATE BLOCKING• STILL IN BLOCKING TAB, CLICK ASSOCIATE AND “ASSOCIATE EDGE TO CURVE”

• CHOOSE EDGE OF BACK BOUNDARY TO ASSOCIATE WITH BACK CURVE WITH CLICK AND

MIDDLE-CLICK UNTIL THE EDGES BECOME GREEN.

• CHOOSE EDGE OF UPPER, FRONT, AND

LOWER BOUNDARIES TO ASSOCIATE WITH UPPER, SEMI-CIRCLE, AND LOWER CURVES WITH

CLICK AND MIDDLE-CLICK UNTIL THE EDGES BECOME GREEN.

CREATE BLOCKING• STILL IN BLOCKING TAB, CLICK ASSOCIATE MENU, CHOOSE “SNAP PROJECT VERTICES”

MODIFY BLOCKING

• MODIFY EXISTING BLOCKING BY EDIT EDGE AND SPLIT EDGE

• SELECT METHOD SPLINE TO EDIT EDGE IN NEAR SEMI-CIRCLE CURVE SO

THAT THE EDGE WILL FOLLOW THE CURVE’S SHAPE.

MODIFY BLOCKING

• AND SELECT METHOD LINEAR TO GRADUALLY CHANGE THE SHAPE OF THE REST EDGE TO

FOLLOW THE CURVES. NOW THE SHAPE OF THE BLOCKING IS THE SAME WITH THE DOMAIN.

MODIFY THE BLOCK• STILL IN BLOCKING MENU, CHOOSE SPLIT BLOCK AND CHOOSE OGRID BLOCK

• SELECT BLOCK BY CLICKING AREA INSIDE OUTER BOUNDARIES, AND SELECT EDGE OF OUTLET

CLICK APPLY

MODIFY THE BLOCK• USE MOVE VERTEX AND CLICK IN LOWER-LEFT WINDOW, THEN MOVE VERTICES SO

THAT THE POSITIONS WILL BE AS IN THE NEXT TWO FIGURE

Select vertices by click

And drag with mouse

To their new places

this vertex

Vertices in their new positions

MODIFY BLOCK

• STILL IN BLOCKING TAB, SELECT SPLIT BLOCK AND CLICK SPLIT BLOCK

MODIFY BLOCK

• SELECT EDGE BEHIND AIRFOIL (THE RED EDGE) BY CLICK, DRAG THE MOUSE TO POSITION OF SPLIT

MODIFY BLOCK

• STILL FROM BLOCKING MENU, SELECT MERGE VERTICES AND COLLAPSE BLOCK ,

MODIFY BLOCK

• ACTIVATE BLOCKS

This edge This block to

collapse

• Click the edge, click block to collapse, middle click

MODIFY THE BLOCK• USE MOVE VERTEX AND CLICK IN LOWER-LEFT WINDOW, THEN MOVE VERTICES SO

THAT THE POSITIONS WILL BE AS IN THE NEXT TWO FIGURE

Select vertices by click

And drag with mouse

To their new places

This vertex

Become at this position

MODIFY BLOCK• STILL FROM BLOCKING TAB, CHOOSE ASSOCIATE , AND ASSOCIATE VERTEX TO POINT

Select this vertex

Click this point

So that the vertex turned to red

MODIFY BLOCK

• STILL FROM BLOCKING TAB, CHOOSE ASSOCIATE , AND ASSOCIATE VERTEX TO POINT

MODIFY BLOCK

• ASSOCIATE TWO VERTICES TO TWO POINTS AS FOLLOWS

This point

This vertex

Vertex turned into red

MODIFY BLOCK• STILL IN BLOCKING TAB, CLICK ASSOCIATE AND “ASSOCIATE EDGE TO CURVE”

Check “project

vertices”

Click this edges to associate with

Airfoil curves, then middle click

Click these two

curves, middle click

MODIFY BLOCKING

• MODIFY EXISTING BLOCKING BY EDIT EDGE AND SPLIT EDGE

• SELECT METHOD LINEAR TO EDIT EDGES INSIDE AIRFOIL TO FOLLOW

AIRFOIL CURVE

Make these edges follow

The airfoil curve

By split edge

With linear method

MODIFY BLOCK

• STILL IN BLOCKING TAB, CHOOSE SPLIT BLOCK , AND CLICK SPLIT BLOCK AND

SELECT THE EDGE (THE RED ONE) BY CLICK SO THAT THERE WILL BE AN ADDITIONAL BLOCK

AROUND AIRFOIL. THEN MIDDLE CLICK TO CONFIRM.

Additional block

Near airfoil

MODIFY BLOCK

• WE WILL DELETE BLOCK INSIDE THE AIRFOIL BY CHOOSING DELETE BLOCK

• CLICK THE BLOCK INSIDE AIRFOIL, MIDDLE CLICK TO CONFIRM TO DELETE.

We will delete this block

• THERE IS NO MORE BLOCK INSIDE THE AIRFOIL

• CHECK BLOCKS IN UPPER LEFT WINDOW TO VIEW BLOCKS IN OUR COMPUTATIONAL DOMAIN

PREPARING FOR MESHING

• IF YOU WORK WITH WORKBENCH, SAVE THE WORKBENCH WITH DIFFERENT NAME BEFORE YOU

DO MESHING

• IF YOU WORK WITH STANDALONE ICEM CFD, SAVE THE PROJECT WITH DIFFERENT NAME BEFORE

YOU DO MESHING

THIS STEP ENSURE YOU TO HAVE A FILE OF GEOMETRY WITH COMPLETE BLOCKING BEFORE MESHING,

SO THAT IF YOU MAKE A MISTAKE IN MESHING YOU CAN GO BACK AND RESTART FROM THIS FILE.

• FOR EXAMPLE, HERE WE SAVE WORKBENCH PROJECT WITH COMPLETE BLOCKING

NOW WE ARE READY FOR MESHING

• FOR EXAMPLE, WE SAVE AN INTERMEDIATE WORKBENCH PROJECT AS FOLLOWS

• YOU CAN DO THE SIMILAR IF YOU WORK WITH STANDALONE ICEM CFD

MESHING• CALCULATING ESTIMATED WALL DISTANCE BY Y+ = 1 (HTTPS://WWW.CFD-

ONLINE.COM/TOOLS/YPLUS.PHP) => CHORD IS 1.O (TO BE ADJUSTED TO 1.625 LATER)

The estimated wall distance

MESHING

• OPEN YOUR AIRFOIL GEOMETRY COMPLETE WITH BLOCKING IN ICEM CFD EITHER FROM

WORKBENCH OR FROM ICEM CFD STANDALONE

• UNCHECK BLOCKS

MESHING

• STILL IN BLOCKING TAB, CHOOSE PRE-MESH PARAM AND EDGE PARAMS

• SELECT EDGE AS IN NEXT SLIDE (RED ONE)

• SET PARAMETER IN LOWER-LEFT WINDOW AS FOLLOWS

This from y+=1

Scroll

MESHING

• SET COPY TO PARAMETERS => ALL PARALLEL EDGES

• CLICK APPLY

• RECOMPUTE PRE-MESH BY RIGHT CLICK ON THE TREE OF “PRE-MESH”

MESHING• WE STILL SEE THAT THE PRE-MESH IS PENETRATING THE AIRFOIL SURFACE

• FROM BLOCKING TAB, CLICK (“GRID PARAMS”) AND CLICK (“UPDATE SIZES”), APPLY

Now, we get pre-mesh that not penetrating airfoil

MESHING

• REPEAT THE PREVIOUS PROCEDURE TO SET MESH AT THE EDGE

• SELECT EDGE AS IN NEXT SLIDE (RED ONE)

• SET PARAMETER IN LOWER-LEFT WINDOW AS FOLLOWS

Scroll This from y+=1

Set copy to parameters => All Parallel Edges

Click apply

• RECOMPUTE PRE-MESH, ACTIVATE PRE-MESH

MESHING

• WE WILL SET MESH FOR OTHER NECESSARY EDGES

MESHING

• SELECT EDGE BEHIND THE AIRFOIL (RED ONE WITH ARROWS)

• SET NODES FOR EXAMPLE 1500

• SET MESH LAW AS “BIGEOMETRIC”

• LET OTHER VALUES AS GIVEN BY SOFTWARE

• SET COPY PARAMETER => ALL PARALLEL EDGES

• CLICK APPLY AND RECOMPUTE PRE-MESH AND ACTIVATE IT

MESHING

• SELECT EDGES OF SURFACE OF AIRFOIL (RED ONE WITH ARROWS)

• CLICK APPLY, RECOMPUTE PRE-MESH AND ACTIVATE IT

MESHING

• SELECT EDGE IN FRONT OF LEADING EDGE OF AIRFOIL(RED ONE WITH ARROWS)

MESHING

• SELECT EDGE NEAR FARFIELD (RED ONE WITH ARROWS)

MESHING

• RIGHT CLICK ON PRE-MESH, AND CHOOSE “CONVERT TO UNSTRUCT MESH”

• IF THERE IS A WINDOW THAT “MESH ALREADY EXISTS”, CLICK “REPLACE”

• NOW WE HAVE ALREADY HAD A MESH

NOW WE ARE READY FOR SIMULATION

SET SOLVER

• CHOOSE OUTPUT MESH TAB AND CHOOSE SELECT SOLVER

• CHOOSE ANSYS FLUENT

SET BOUNDARY CONDITIONS• FROM OUTPUT MESH TAB, CHOOSE (“BOUNDARY CONDITIONS”)

• FIRST, SET BOUNDARY CONDITION FOR “AIRFOIL” THAT APPEARS IN “EDGES”

• FOLLOW STEPS IN THE NEXT SLIDE

1. Click “create new”

2. Choose “wall”

3. Click “Okay”

SET BOUNDARY CONDITIONS

• THEN SET BOUNDARY CONDITION OF FARFIELD

• FROM “?MIXED/UNKNOWN” , CHOOSE FARFIELD

2. Choose “velocity inlet”

3. Click “okay”

• SET THE BOUNDARY CONDITION OF FLUID

• FROM “? MIXED/UNKNOWN”, CHOOSE “FLUID”

2. Choose “fluid”

3. Click “Okay”

• THEN WE WILL HAVE COMPLETE BOUNDARY CONDITIONS

• CLICK ACCEPT

WRITE OUTPUT FILE• FROM OUTPUT MESH TAB, CHOOSE (“WRITE OUTPUT”)

WRITE OUTPUT FILE

• CLICK OPEN FOR DEFAULT *.UNS FILE WHEN ASKED

WRITE OUTPUT FILE

• BECAUSE WE WILL OPEN FLUENT MESH BY FLUENT STANDALONE, NAME MESH WITH A UNIQUE

• MAKE SURE THE CASE IS “2D”

Remember this working directory and the name of mesh file

PREPARING FOR SIMULATION IN FLUENT

• SAVE YOUR PROJECT AND CLOSE YOUR WORKBENCH OR STANDALONE ICEM CFD

• LAUNCH STANDALONE FLUENT IN 2D CASE WITH DOUBLE PRECISION OPTION

• CLICK “OK”

SIMULATION IN FLUENT

• CHOOSE FILE => READ => MESH

• CHOOSE THE MESH FILE

THAT WE SAVED

PREVIOUSLY

• CLICK OK

• MESH DISPLAYED IN FLUENT

• FIRST, WE WILL SCALE THE GEOMETRY AND MESH OF AIRFOIL TO HAVE A CHORD LENGTH OF

1.625.

• FROM GENERAL SETUP, CLICK “SCALE”

SIMULATION IN FLUENT• IN SCALING MENU, CHOOSE “SPECIFY SCALING FACTOR”

• ENTER FACTORS OF 1.625 FOR BOTH X AND Y, CLICK “SCALE” UNTIL THE VALUES ARE

CHANGED

• AFTER SCALING, AGAIN IN GENERAL SETUP, WE MAKE SURE THAT THE SETUP IS AS BELOW

• IN MODELS SETUP, CLICK “VISCOUS” AND CHOOSE “SPALART ALMARAS”

• CLICK OK

• IN MATERIALS SETUP, CHOOSE UNDER “FLUID”: “AIR” UNTIL A POP WINDOW APPEARS

• MAKE SURE THAT WE CHOOSE AIR AT SEA LEVEL CONDITION AS IN THE NEXT SLIDE

• CLICK CHANGE/CREATE AND CLICK CLOSE

SIMULATION IN FLUENT• IN CELL-ZONE CONDITIONS, CLICK “FLUID”, AND CLICK “OPERATING CONDITIONS”

SIMULATION IN FLUENT• ENTER 1 ATM OR 101325 PA FOR A POSITION OF FARFIELD IN THE FRONT. HERE WE USE THE

POINT OF (-32.5,0)

• WE CHECK BOUNDARY CONDITIONS

• AIRFOIL IS WALL

• FARFIELD IS VELOCITY INLET

• CLICK “EDIT”

SIMULATION IN FLUENT• SET VELOCITY AS “COMPONENT” AND ENTER THE VALUE OF FREESTREAM VELOCITY OF 26.82

• INT_FLUID AS “INTERIOR”

• SET REFERENCE VALUES AS FOLLOWS

SIMULATION IN FLUENT• IN MONITORS, WE WILL MAKE ADDITIONAL MONITORS OF CL, CD, AND CM. CLICK “REPORT

PLOTS” AND CLICK “NEW”

SIMULATION IN FLUENT• CREATE DRAG MONITORS

• DRAG REPORT DEFINITION

• CREATE LIFT MONITORS

• LIFT REPORT DEFINITIONS

• CREATE MOMENT MONITORS

• MOMENT REPORTS DEFINITION0.25chord

• IN SOLUTION INITIALIZATION, CHOOSE “STANDARD INITIALIZATION” AND COMPUTE FROM

“FARFIELD” AND LET OTHER VALUES AUTOMATICALLY DETERMINED BY SOFTWARE

• CLICK “INITIALIZE”

Click “Initialize”

• NOW, WE ARE READY TO RUN CALCULATION.

• ENTER ITERATION OF 500, WHICH COULD BE REPEATED LATER IF NOT CONVERGENT.

• BECAUSE NOT SET, THE CONVERGENCE CRITERIA WILL BE BY DEFAULT 1E-4.

• CLICK “CALCULATE” TO START SIMULATION IN FLUENT

• SIMULATION WILL TAKE ABOUT 8800 - 9800 ITERATIONS FOR ABOUT 7-16 HOURS

SIMULATION TO CONVERGED (AT I7 OR I5 PROCESSOR).

• YOU CAN’T DO ITERATION FOR EACH 1000 ITERATIONS AND SAVE THE CASE & DATA FOR

EACH 1000 ITERATIONS AND CONTINUE (DO NOT INITIALIZE, JUST CLICK CALCULATE)

• DON’T FORGET TO SAVE THE WORK FROM FILE => WRITE => CASE & DATA

• THE RESULT WILL BE EVALUATED BY EXPERIMENTAL RESULT AFTER POST-PROCESSING

NOW WE ARE READY FOR POST-PROCESSING

DISPLAYING LIFT COEFFICIENT

• WE WILL DISPLAY THE RESULT OF LIFT. CLICK “REPORTS”, CLICK “FORCES”, CLICK “SET UP” AND

A POP-UP WINDOW APPEARS

• THEN SET “DIRECTION VECTOR” OF Y AS 1 BECAUSE FOR ANGLE OF ATTACK 0 DEGREE, THE

LIFT FORCE IS IN Y POSITIVE DIRECTION.

• CLICK “PRINT” AND A LIFT COEFFICIENT VALUE WILL APPEAR IN LOWER-RIGHT WINDOW

DISPLAYING DRAG COEFFICIENT

• WE WILL DISPLAY THE RESULT OF DRAG. CLICK “REPORTS”, CLICK “FORCES”, CLICK “SET UP”

AND A POP-UP WINDOW APPEARS

• THEN SET “DIRECTION VECTOR” OF X AS 1 BECAUSE FOR ANGLE OF ATTACK 0 DEGREE, THE

DRAG FORCE IS IN X POSITIVE DIRECTION.

• CLICK “PRINT” AND YOU WILL SEE THE DRAG COEFFICIENT VALUE IN LOWER-RIGHT WINDOW

WRITING THE LIFT AND DRAG COEFFICIENT TO FILE• YOU CAN SAVE THE LIFT OR DRAG COEFFICIENT TO FILE BY CLICK “WRITE” INSTEAD OF “PRINT”

THEN INPUT THE FILE NAME

COMPARING WITH EXPERIMENTAL DATA

• IT IS IMPORTANT TO COMPARE OUR RESULT OF CFD SIMULATION WITH ANOTHER DATA TO

KNOW THE QUALITY OF RESULT. HERE WE WILL COMPARE WITH EXPERIMENTAL DATA OF

AIRFOIL NACA 65(2)-415 TAKEN FROM NACA REPORT. THE FILE OF NACA REPORT WILL BE

PROVIDED.

From experimental at 0 degree with Reynolds number 3e6, lift coefficient is 0.3 and drag coefficient is 0.007

COMPARING WITH EXPERIMENTAL DATA

• FROM EXPERIMENTAL DATA, AT 0 DEGREE WITH REYNOLDS NUMBER OF 3 MILLIONS, LIFT

COEFFICIENT IS 0.3 AND DRAG COEFFICIENT IS 0.007 (CL/CD (AERODYNAMICS EFFICIENCY) IS

ABOUT 42)

• THE RESULT OF CFD SIMULATION, THE LIFT COEFFICIENT IS 0.548 AND DRAG COEFFICIENT IS

0.0128 (CL/CD (AERODYNAMICS EFFICIENCY) IS ABOUT 42)

• WE CAN SEE THAT THE CFD SIMULATION IS IN MEDIUM QUALITY, BECAUSE IT HAS THE SAME

ORDER WITH EXPERIMENTAL RESULT, BUT IF YOU WILL USE CFD SIMULATION RESULT FOR A

PROJECT, YOU SHOULD MODIFY THE MESHING OR CHANGE TURBULENCE MODEL USED.

DISPLAYING ABSOLUTE PRESSURE CONTOUR

• IT IS IMPORTANT TOO TO CHECK WHETHER THE TOTAL PRESSURE CONTOUR IS AS IT SHOULD

BE, WHICH IS 1 ATM AT FARFIELD.

• CLICK “GRAPHICS”, CLICK “CONTOURS”, CLICK “SET UP” UNTIL A POP-UP WINDOW APPEARS.

• CHOOSE “PRESSURE” AND “ABSOLUTE PRESSURE”

• CHOOSE “INT_FLUID” OF SURFACE

• CLICK “COMPUTE”, WAIT, AND CLICK “SAVE/DISPLAY”.

• USE “FIT TO WINDOW” TO SEE THE WHOLE DOMAIN.

The CFD simulation results in correct values of 1 atm (about 1.01325e5 Pa) at farfield

SHOWING STATIC PRESSURE CONTOURS• WE WILL CHECK STATIC PRESSURE CONTOURS AROUND THE AIRFOIL WHETHER IT HAVE A

NEGATIVE STATIC PRESSURE ON THE UPPER SURFACE AS IT SHOULD BE.

• CLICK “GRAPHICS”, CLICK “CONTOURS”, CLICK “SET UP” UNTIL A POP WINDOW APPEARS.

• CHOOSE “PRESSURE” AND “STATIC PRESSURE”

• CHOOSE “INT_FLUID” OF SURFACE

• CLICK “COMPUTE”, WAIT, AND CLICK “SAVE/DISPLAY”. CHOOSE THE WINDOW TO VIEW THE

STATIC PRESSURE CONTOUR.

• USE “ZOOM TO AREA” TO SEE CONTOUR NEAR THE AIRFOIL.

SAVING CONTOUR PLOT

• IF YOU NEED TO SAVE THE CONTOUR PLOT, CHOOSE “SAVE PICTURE” FROM THE ICON.

• THEN A POP UP WINDOW WILL APPEAR, ENTER THE THE FILE TYPE.

• CLICK “SAVE”, AND ENTER THE FILE NAME.

EXAMPLE OF JPEG PICTURE OF STATIC PRESSURE CONTOUR

EVALUATION OF STATIC PRESSURE CONTOUR

• WE SEE FROM STATIC PRESSURE CONTOUR THAT THERE IS A NEGATIVE STATIC PRESSURE AREA

ON UPPER SURFACE OF AIRFOIL, THIS FACT CONFIRM THAT THE CFD SIMULATION IS CORRECT

ACCORDING TO THEORY.

DISPLAYING VELOCITY CONTOUR

• WE WILL CHECK THAT AT FARFIELD, THE VELOCITY SHOULD BE 26.82 M/S.

• CLICK “GRAPHICS”, CLICK “CONTOURS”, CLICK “SET UP” UNTIL A POP WINDOW APPEARS.

• CHOOSE “VELOCITY”, CHOOSE “VELOCITY MAGNITUDE”

• CLICK “COMPUTE”, WAIT, CLICK “SAVE/DISPLAY”

• SELECT THE WINDOW TO SHOW THE CONTOUR OF VELOCITY

• SELECT “FIT TO WINDOW” TO SEE WHOLE OF DOMAIN

We see that the velocity near of farfield is of correct value of 26.82 m/s

NOTE ON POST-PROCESSING

• THERE IS MANY THINGS CAN BE DISPLAYED IN POST-PROCESSING, YOU CAN TRY BY YOURSELF

• FOR EXAMPLE, DISPLAYING VELOCITY VECTORS NEAR AIRFOIL SURFACE, AND CHECKING

VELOCITY PROFILE OF BOUNDARY LAYER.

Activity sheet airfoil and wing fluent simulation

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