AeroWindTunnel 6.0
Instruction Manual

Airplane Flight Dynamics & Stability Analysis
for Gliding and Powered Flight


Purchase

| MAIN PAGE | PRODUCTS | CONSULTING | MISSION | RESUME |
Copyright 1999-2013 John Cipolla/AeroRocket
You are "New" Visitor  0001512

AeroWindTunnel is a Microsoft Windows, slider-input based computer program to determine if the criteria for longitudinal (pitch) or up and down and lateral (yaw) or side to side stability are satisfied for an airplane or glider in free flight. The necessary criteria for longitudinal static stability are that CM_0 the moment coefficient around the center of gravity at zero lift must be positive and CMcg_a the slope of the moment coefficient (CMcg) verses airplane angle of attack (a) curve must be negative. The aerodynamics routines within AeroWindTunnel estimate subsonic and supersonic parasite (zero-lift) drag coefficient by the component build-up method using a well-known flat-plate skin friction drag coefficient formula that is a function of Reynolds number and Mach number. Base drag and separation drag, known percentages or functions of skin friction drag according to Hoerner and others, leads to  an "equivalent skin friction drag coefficient" that is a function of skin friction and separation drag effects.

In addition, AeroWindTunnel models airframe plan-view and side-view as a collection of 19 conical frustrums or trapezoids to determine center of pressure (Xcp), drag (CD), lift (CL), lift slope (Cna) and moment slope (Cma) coefficients using an "exact" mathematical approach. The theory for determining the aerodynamic properties of an airframe in AeroWindTunnel can be found in THE THEORETICAL PREDICTION OF THE CENTER OF PRESSURE by James and Judith Barrowman. AeroWindTunnel uses the resulting Barrowman equations for a conical frustrum as described in Barrowman's paper for plan-view and side-view dimensions to determine airframe aerodynamics. AeroWindTunnel integrates 19 conical frustrums to determine the aerodynamics of almost any complex airframe. The approach is mathematically exact as the number of frustrum elements is increased. However, 19 conical frustrums provide excellent accuracy for most complex aerodynamic shapes. Therefore, panel methods, complex CFD analyses and DATCOM are avoided with their inherent problems. Instead, AeroWindTunnel uses an exact mathematical technique that has demonstrated an excellent history of success.

Wave drag for supersonic flow is rapidly determined by an empirical Sears-Haack method that is a correlation of the actual Sears-Haack wave drag experienced by aircraft in supersonic flight. Using these methods zero-lift drag, drag due to lift, lift curve, lift slope, moment curve, and moment slope are quickly determined and have been validated from 0.3 Mach to 3.5 Mach using the AeroRocket wind tunnel and other sources. AeroWindTunnel is not a CAD program for determining weights and balance of aircraft and gliders but instead is a computer-based, conceptual-design wind tunnel program that uses slider-bar entry and imported fuselage shapes to quickly estimate stability of airplanes and gliders. Other features include display of elevator angle to trim given aircraft weight and velocity. Also, velocity and thrust required for level flight given aircraft weight are quickly determined.

AeroWindTunnel is AeroRocket's first true wind tunnel simulation program for determining model airplane and glider stability and whether or not a glider or model airplane is stable enough to fly. AeroWindTunnel is a true wind tunnel simulation program in that it determines whether an airplane is stable based on the position of the wing-body neutral point relative to the center of gravity and whether the relative contribution of the wings, fuselage and vertical fins to the forces and moments around the airplane center of gravity are sufficient for longitudinal and lateral stability. To perform a flight dynamics analysis AeroWindTunnel predicts CL, CD, CM and the derivatives of these coefficients relative to angle of attack as a function of the absolute and geometric angle of attack of the airplane or rocket in flight. AeroWindTunnel includes a bonus program called 2D-WING for the aerodynamic analysis of two-dimensional and finite aspect ratio (AR) airfoil sections.

This Instruction Manual for AeroWindTunnel uses the HL-20 and X-30 Space Planes displayed below to illustrate the usefulness of AeroWindTunnel for determining the flight characteristics of realistic, real-world flight vehicles. In addition, an F-16 Type Jet Airplane has been included to briefly describe AeroWindTunnel's ability for the analysis of supersonic jet airplanes.
Also, a new model airplane design called AeroEagle has been included to illustrate the usefulness of AeroWindTunnel for determining glider and model airplane aerodynamics. Please see the present collection of aircraft and space plane projects analyzed using AeroWindTunnel. Finally, please read our disclaimer for this computer program.

Airplane Flight Dynamics Theory
Static stability and control for roll, pitch and yaw are necessary criteria for the design of conventional airplanes and gliders. Assume a rigid airplane or glider with elevator fixed in some position is tested in a wind tunnel or analyzed using CFD to determine its variation of Cm (pitch-moment coefficient) around the airplane center of gravity (cg) verses absolute angle of attack (aa). Figure-1, Cm vs. absolute angle of attack.The relationship between Cm verses absolute angle of attack is illustrated in Figure-1. The value of Cm around the center of gravity where lift and absolute angle of attack are zero is called Cm0. Further, the angle of attack where Cm = 0 is called the equilibrium point or trim angle of attack. Also, the slope of the pitch-moment coefficient (Cm) verses angle of attack curve is Cma whose sign determines whether or not an airplane or glider is longitudinally stable. The primary condition for pitch equilibrium is that the pitch-moment coefficient (Cm) around the center of gravity must be zero. An airplane is considered longitudinally stable if the slope of the pitch-moment verses angle of attack curve (Cma) is negative as illustrated by curve-b. For curve-b, the incremental moment (DCm) following a disturbance (like a gust of wind) is negative which tends to restore a to its equilibrium value, ae. In contrast, for positive Cma as in curve-a the opposite effect occurs where a positive incremental change in moment (DCm) tends to increase the angle of attack making the airplane unstable in pitch. For curve-b the situation is similar to that of a mass on a spring where aerodynamic and mechanical disturbances tend to restore a to the equilibrium value represented by ae The aerodynamic property that tends to hold a constant at its equilibrium value is called "static stability". In other words an airplane exhibits "positive pitch-stability" when curve-b represents the response of an airplane or glider to changes in angle of attack. Positive pitch stiffness (Cma < 0) is an important design criterion which makes its measurement or computation critical. A complete description of airplane stability is beyond the scope of this discussion but more information can be found in the book Introduction to Flight, by John D. Anderson. Please refer to the AeroRocket list of references upon which the theory for AeroWindTunnel and many other AeroRocket software are based or the full list of AeroRocket's reference books.

SUMMARY OF PITCH STABILITY CRITERIA
1) Cm
a (slope of the pitch-moment verses angle of attack curve) must be negative.
2) Cm0 (value of Cm around the center of gravity where the lift is zero) must be positive for
a > 0.
3) The equilibrium angle of attack (
ae) must fall within the normal flight range of angle of attack for the airplane or glider.



HL-20, X-30 NASP and  Supersonic Jet Airplane designs used to illustrate AeroWindTunnel
Click models for more information
 

Basic Analysis Procedure
Please Use One of Two Methods

METHOD-1: SLIDER-BAR
Define all aerodynamic surfaces like wing span and wing chord by selecting the Slider inputs option button. Then, by selecting either the TAIL/ELEVATOR or ELEVATOR option button the user can define an airplane composed of main wing and tail or a tail-less airplane. The input data required for slider-bar entry include number of vertical fins, wing dihedral angle, main wing setting angle,  zero-lift aoa (angle of attack), wing-body downwash angle, wing-body downwash gradient, wing-body moment coefficient, elevator effectiveness lift-slope, and tailless elevator moment. Further, by selecting the Import fuselage option button the user can replace the simple slider-bar symmetric fuselage with a more complex non-symmetric fuselage. If the imported fuselage option is selected, click the Fuselage-Aerodynamics command in the main menu. When the Fuselage Geometry screen appears click Input-Airframe-Data to import the fuselage geometry and then enter the other inputs on this screen to complete the fuselage analysis. Where, wave drag for wings having the following cross-sections are specified as: Single wedge (KLE=1), Symmetrical double wedge (KLE=4), Biconvex section (KLE=5.3, Streamline foil with x/c=50% (KLE=5.5), Round-nose foil with x/c=30% (KLE=6.0), Slender elliptical airfoil section (KLE=6.5) and Double wedge with maximum thickness at arbitrary x/c location (KLE=[c/x]/[1-x/c]).
 
Defining the geometry of the airplane is performed by manipulating the slider-bars corresponding to MAIN WING, TAIL/ELEVATOR or ELEVATOR and VERTICAL FIN. Then, in the RESULTS AND LEVEL FLIGHT ANALYSIS data entry section the user defines the geometric angle of attack, elevator deflection (if required), airplane velocity and airplane weight. A complete analysis is performed each time data is modified and if the airplane will fly the statement, SUCCESSFUL ANALYSIS - AIRPLANE WILL FLY appears in the comment box. If an unsuccessful analysis occurs reasons for the failure appear in the FLIGHT RESULTS portion of the PLAN-VIEW AND SIDE-VIEW PLOTS data entry section. When those conditions are satisfied the airplane will fly.

Please note
: For AeroWindTunnel wing span is the total distance from wing-tip to wing-tip including fuselage diameter separation between wing-roots. In addition, tail span is the distance from tail-tip to tail-tip where fuselage diameter separation is not included in span length. Finally, fin span is the total distance from fin-root to fin-tip where fuselage diameter separation is not included in span length.
 
METHOD-2: MANUAL ENTRY
Define all aerodynamic surfaces and flight coefficients for an airplane composed of wing and tail with elevator by selecting the Manual inputs option button in the
PLAN-VIEW AND SIDE-VIEW PLOTS data entry section. The input data required for manual input include main wing exposed surface area, main wing mean aerodynamic chord, horizontal tail exposed surface area, vertical fin exposed surface area, number of vertical fins, wing-body cg location from wing LE (Leading Edge), distance from airplane cg (center of gravity) to tail's ac (aerodynamic center), airframe fineness ratio (Lmax/Dmax), wing dihedral angle, main wing setting angle, tail dihedral angle, horizontal tail setting angle, wing-body ac from wing LE, zero-lift aoa, wing-body downwash angle, wing-body downwash gradient, wing-body moment coefficient, wing-body lift-slope coefficient, horizontal tail lift-slope coefficient, elevator effectiveness lift-slope, and tailless elevator moment. Then, in the RESULTS AND LEVEL FLIGHT ANALYSIS data entry section the user defines the geometric angle of attack, elevator deflection (if required), airplane velocity and airplane weight. A complete analysis is performed each time data is modified and if the airplane will fly the statement, SUCCESSFUL ANALYSIS - AIRPLANE WILL FLY appears in the comment box. If an unsuccessful analysis occurs reasons for the failure appear in the FLIGHT RESULTS portion of the PLAN-VIEW AND SIDE-VIEW PLOTS data entry section. When those conditions are satisfied the airplane will fly.


AeroWindTunnel Step-by-Step Examples
1) MACH 0.30 HL-20 SPACEPLANE (Slider-Input) Back
To run the example of the HL-20 operating at 0.3 Mach (below) using the provided input files please click the project file, HL-20 PROJECT_30deg.DAT using the main screen File menu. Then, when prompted click HL-20 FUSELAGE.TXT from the Fuselage Geometry and Wing-Fuselage Aerodynamic Center (ac) Location screen. Close this screen when ready. To see coefficient plots click Plot-Coefficients on the main screen to display the Coefficient Plots screen. Do not forget to select the reference area by clicking Body Planform Area in the Plot-Reference-Area command located in the top menu of the Coefficient Plots screen. Then, click Plot-Experimental-Data in the top menu. In the File menu, click Load Experimental Data and load HL-20_WIND_TUNNEL_DATA_30deg.EXP, the wind tunnel experimental data from the paper, Aerodynamic Characteristics of the HL-20 at M = 3.5. Finally, click PLOT and then CLOSE to see the plots.

FILES REQUIRED
HL-20 PROJECT_30deg.DAT (Project file)
HL-20 FUSELAGE.TXT (Fuselage geometry file)
HL-20_WIND_TUNNEL_DATA_30deg.EXP (Experimental data file)

FUSELAGE GEOMETRY IMPORT FILE FORMAT (21 lines of formatted data)
Line-1: Fuselage length from tip to end in appropriate units.
Line-2: Station-1 plan-view width, Station-1 height above side-view centerline, Station-1 height below side-view centerline.
...
Line-21: Station-20 plan-view width, Station-20 height above side-view centerline, Station-20 height below side-view centerline.

Please click Fuselage-Aerodynamics for HL-20 fuselage definition and Plot-Coefficients to see coefficient plots and validation with HL-20 data.
Figure-2, AeroWindTunnel main screen used to define the basic  geometry of the HL-20 spacecraft.

STEP-BY-STEP INSTRUCTIONS
HL-20 (M = 0.30, Elevator Deflection = -30 degrees and 0.0 degrees)

1) SET UNITS AND PLOT AREA DIMENSIONS

Click Units located in the top menu and check MKS to set units in meters.
Click Max X-Y located in the top menu to set the model's Maximum X and Y dimension to 10x10 meters.

2) WING DIMENSIONS
Using MAIN WING Span slider bar set wing span length (exposed) to 4.73 meters.
Using MAIN WING Root slider bar set wing root length to 2.93 meters.
Using MAIN WING Tip slider bar set tip length to 1.21 meters
Using MAIN WING Tip Sweep distance slider bar set tip sweep distance to 2.55 meters.
Using MAIN WING LE location from nose tip slider bar set wing LE location from nose-tip to
5.11 meters.

3) HORIZONTAL TAIL DIMENSIONS
Click the ELEVATOR option button to indicate the airplane is tailless and uses an elevator for pitch control.
Because the airplane is tailless, set all five tail slider bars to 0.0 by pulling each slider to the far left.

4) VERTICAL FIN (S) DIMENSION
Set WING, TAIL and ELEVATOR efficiency (
h) to 100% using the three vertical slider bars.
Using VERTICAL FIN Span slider bar set vertical fin span to 0.725 meters.
Using VERTICAL FIN Root slider bar set fin root length to 1.62 meters.
Using VERTICAL FIN Tip slider bar set tip distance to 0.0 meters.
Using VERTICAL FIN Tip Sweep distance slider bar set tip sweep distance to 1.62 meters.
Using VERTICAL FIN LE Location from nose tip slider bar
set fin LE location from nose-tip = 6.41 meters.

5) FUSELAGE GEOMETRY AND FLOW CHARACTERISTICS
Check the Import Fuselage box to activate the Fuselage Aerodynamics command in the top menu.
Click the activated Fuselage Aerodynamics command located in the top menu to bring up the Fuselage Geometry ... screen.
Click Input Airframe Data located in the top menu to import HL-20 FUSELAGE.TXT, the geometry file for the HL-20 fuselage.
Unclick Boat tail effect on base drag for the HL-20 fuselage.
Close the Fuselage Geometry ... Input Screen.

FUSELAGE GEOMETRY IMPORT FILE FORMAT (21 lines of formatted data)
Line-1: Fuselage length from tip to end in appropriate units.
Line-2: Station-1 plan-view width, Station-1 height above side-view centerline, Station-1 height below side-view centerline.
...
Line-21: Station-20 plan-view width, Station-20 height above side-view centerline, Station-20 height below side-view centerline.

6) WING AND ELEVATOR COEFFICIENTS
On the main screen in the Manual Entry section enter the following.
Flight altitude above sea level = 0.0 meters
Number of vertical fins = 0.
Wing dihedral angle = 50 degrees .
Main wing setting angle = -6.6 degrees.
Zero-lift angle of attack (aoa) = 2.5 degrees.
Wing-body down-wash angle = 0.0 degrees.
Wing-body down-wash gradient = 0.0
Wing-body moment coefficient = 0.0
Elevator effectiveness lift slope = 0.00912 1/deg
Tailless elevator moment = -0.00348 1/deg

7) LEVEL FLIGHT REQUIREMENTS
Using the CG location from wing LE slider bar set the CG location to -0.65 meters from the wing LE or 0.54 L from the nose-tip.
In the Airplane velocity (V2) text box set airplane velocity to 103.177 meters/sec (Mach = 0.3032)
In the Total airplane weight in level trim flight text box set airplane weight to 10884 kg.
Set the Geometric angle of attack using the up-down command to 10 degrees.
Using the Elevator deflection text box set the elevator deflection to -30 degrees.

8) PLOT RESULTS
Plot the results by clicking the Plot-Coefficients command.
Select Body Planform area as the reference area for CD, CL, CM etc. by using the Plot-Reference-Area command in the top menu.
Use the Plot-Experimental-Data command to plot the coefficient data by bringing up the Insert Experimental Data Into Coefficient Plots Screen.
Under File load the data to be plotted by clicking the Load Experimental Data command.

In AeroWindTunnel simply click Input Airframe Data to import the x-y dimensions of the HL-20 plan-view and side-view fuselage shape.

Figure-3, AeroWindTunnel Fuselage geometry import screen. Simply define 20 plan-view and side-view station locations.

HL-20 at M = 0.30 and Elevator Deflection at -30 degrees
Please click Plot-Experimental Data to see the screen containing HL-20 wind tunnel data to be compared with AeroWindTunnel plot results.

Figure-4a, AeroWindTunnel results compared with HL-20 wind tunnel test data from the paper, Aerodynamic Characteristics of the HL-20 at M = 0.3 and elevator set to -30 degrees. Blue circle-lines represent AeroWindTunnel data for the HL-20 and the red circle-lines represent results from the paper, referenced above.
 

HL-20 at M = 0.30 and Elevator Deflection re-set to 0.0 degrees

Figure-4b, AeroWindTunnel results compared with HL-20 wind tunnel test data from the paper, Aerodynamic Characteristics of the HL-20 at M = 0.3 and elevator set to 0.0 degrees. Blue circle-lines represent AeroWindTunnel data for the HL-20 and the red circle-lines represent results from the paper, referenced above.
 

1a) MACH 3.5 HL-20 SPACEPLANE
To run the example of the HL-20 operating at 3.5 Mach (below) using the provided input files please click the project file, HL-20 PROJECT_30deg_M35.DAT using the main screen File menu. Then, when prompted click HL-20 FUSELAGE.TXT from the Fuselage Geometry and Wing-Fuselage Aerodynamic Center (ac) Location screen. Close this screen when ready. To view coefficient plots click Plot-Coefficients on the main screen to display the Coefficient Plots screen. Do not forget to select the reference area by clicking Body Planform Area in the Plot-Reference-Area command located in the top menu of the Coefficient Plots screen. Then, click Plot-Experimental-Data in the top menu. In the File menu, click Load Experimental Data and load HL-20_WIND_TUNNEL_DATA_30degM35.EXP, the wind tunnel experimental data from the paper, Aerodynamic Characteristics of the HL-20 at M = 3.5. Finally, click PLOT and then CLOSE to see the plots.

FILES REQUIRED
HL-20 PROJECT_30deg_M35.DAT (Project file)
HL-20 FUSELAGE.TXT (Fuselage geometry file)
HL-20_WIND_TUNNEL_DATA_30degM35.EXP (Experimental data file)

Figure-5, AeroWindTunnel results compared with HL-20 wind tunnel test data from the paper, Aerodynamic Characteristics of the HL-20 at M = 3.5. Blue circle-lines represent AeroWindTunnel data for the HL-20 and the red circle-lines represent results from the paper, referenced above.

 

2) X-30 SPACEPLANE (Slider-Input) Back
To run the example of the X-30 spaceplane flying at 100 meters/sec (below) using the provided input files please click the project file, X-30 PROJECT.DAT using the main screen File menu. Then, when prompted click X-30 FUSELAGE.TXT from the Fuselage Geometry and Wing-Fuselage Aerodynamic Center (ac) Location screen. Close this screen when ready. To see coefficient plots click Plot-Coefficients on the main screen to display the Coefficient Plots screen. Do not forget to select the reference area by clicking Exposed-Wing-Area in the Plot-Reference-Area command located in the top menu of the Coefficient Plots screen. Then, click Plot-Experimental-Data in the top menu. In the File menu, click Load Experimental Data and load X-30_WIND_TUNNEL_DATA.EXP, the wind tunnel experimental results obtained using the subsonic AeroRocket wind tunnel. Finally, click PLOT and then CLOSE to see the plots.

FILES REQUIRED
X-30 PROJECT.DAT (Project file)
X-3- FUSELAGE.TXT (Fuselage geometry file)
X-30_WIND_TUNNEL_DATA.EXP (Experimental data file)

Please click Fuselage-Aerodynamics for X-30 fuselage definition and Plot-Coefficients to see coefficient plots and validation with X-30 data.

Figure-6, AeroWindTunnel main screen used to define the basic  geometry of the X-30 spaceplane.

In AeroWindTunnel simply click Input Airframe Data to import the x-y dimensions of the X-30 plan-view and side-view fuselage shape.

Figure-7, AeroWindTunnel Fuselage geometry import screen. Simply define 20 plan-view and side-view station locations.



3-DIMENSIONAL ORTHOGRAPHIC DISPLAY OF AIRCRAFT MODEL GEOMETRY
SHADING AND HIDDEN LINE CONTROLS COMING IN A FUTURE RELEASE

Figure-8, Three-dimensional wireframe display of aircraft model geometry illustrating rotation, translation and magnification controls. Access the 3D Plot controls by clicking 3D-Plots and then Show in the top menu. Rotate the model around the X, Y and Z axes by simply pulling the slider bar controls for rotations ranging from 0.0 to 360 degrees. Then, magnify and translate the model into position using the lower set of slider bar controls. Rapid response and small program size is achieved by using native code algorithms not canned programs like OpenGL used by the leading rocket simulation programs. More complex operations like hidden lines, shading and textures coming in future upgrades.

Please click Plot-Experimental Data to see the screen containing X-30 AeroRocket subsonic wind tunnel data compared with AeroWindTunnel plot results.

Figure-9, AeroWindTunnel results compared with X-30 AeroRocket subsonic wind tunnel test data. Blue circle-lines represent AeroWindTunnel data for the X-30 and the red circle-lines represent results from testing the X-30 in the AeroRocket subsonic wind tunnel. The X-30 is pictured in Figure-10 being tested in the AeroRocket wind tunnel.


Figure-10, X-30 being tested in the AeroRocket wind tunnel.



 
3) AIRPLANE FROM INTRODUCTION TO FLIGHT (Slider-Input)
Please see examples 7.3, 7.4, 7.5, 7.6, 7.7 and 7.8 from "Introduction to Flight" by John D. Anderson

FILE REQUIRED
GLIDER_PROJECT_INTRO_TO_FLIGHT_LD10_SLIDER-2.DAT (Project file)


Figure-11, AeroWindTunnel main screen used to define the basic  geometry of the examples in Introduction to Flight.

STEP-BY-STEP INSTRUCTIONS
Airplane from "Introduction to Flight" by John D. Anderson, Examples 7.3, 7.4, 7.5, 7.6, 7.7 and 7.8
1) SET UNITS PLOT AREA DIMENSIONS AND DEFINE FUSELAGE
Click Units located in the top menu and check MKS for units in meters.
Click Max X-Y located in the top menu to set model's Maximum X and Y dimension to 1x1 meters.
Using Airframe/Fuselage Maximum length slider bar set maximum fuselage length to 0.48 meters.
Using Airframe/Fuselage Maximum diameter slider bar set maximum fuselage diameter to 0.048 meters.
Using CG location from wing LE slider bar set the CG location to 0.035 meters from the wing LE.

2) WING DIMENSIONS
Using MAIN WING Span slider bar set wing span to 1.0 meters.
Using MAIN WING Root slider bar set wing root length to 0.1 meters.
Using MAIN WING Tip slider bar set tip length to 0.1 meters
Using MAIN WING Tip sweep distance slider bar set tip sweep distance to 0.0 meters.
Using MAIN WING LE location from nose tip slider bar set wing LE location from nose-tip to 0.202 meters.

3) HORIZONTAL TAIL DIMENSIONS
Set WING, TAIL and ELEVATOR efficiency (h) to 77.9%,  129.2% and 100.0% using vertical slider bars.
Using TAIL Span slider bar set tail span to 0.286 meters.
Using TAIL Root slider bar set fin root length to 0.07 meters.
Using TAIL FIN Tip slider bar set the tip distance to 0.07 meters.
Using TAIL FIN Tip Sweep distance slider bar set the tip sweep distance to 0.0 meters.
Using TAIL LE Location from nose tip slider bar set the LE location to 0.39 meters.

4) VERTICAL FIN (S) DIMENSION
Using VERTICAL FIN Span slider bar set vertical fin span to 0.1 meters.
Using VERTICAL FIN Root slider bar set fin root length to 0.05 meters.
Using VERTICAL FIN Tip slider bar set tip distance to 0.05 meters.
Using VERTICAL FIN Tip Sweep distance slider bar set the tip sweep distance to 0.0 meters.
Using VERTICAL FIN LE Location from nose tip slider bar set the LE location from nose-tip to 0.428 meters.

5) WING AND ELEVATOR COEFFICIENTS
On the main screen in the Manual Entry section enter  the following.
Number of vertical fins =1.
Flight altitude above sea level = 0.0 meters
Wing dihedral angle = 0.0 degrees .
Main wing setting angle = 0.0 degrees.
Tail dihedral angle = 0.0 degrees.
Horizontal tail setting angle = 2.7 degrees
Zero-lift aoa = -1.5 degrees.
Wing-body down wash angle = 0.0 degrees.
Wing-body down-wash gradient = 0.35
Wing-body moment coefficient = -0.032
Elevator effectiveness lift slope= 0.04

6) LEVEL FLIGHT REQUIREMENTS
In Airplane velocity (V2) text box set airplane velocity to 61 meters/sec (Mach = 0.17736)
In Total airplane weight in level, trim flight text box set airplane weight to 12.183 kg.
Set Geometric angle of attack using the up-down command to 7.88 degrees.
Set elevator deflection in the Elevator deflection text box to 0.0 degrees.

7) PLOT RESULTS
Plot the results by clicking the Plot-Coefficients command.
Select Exposed wing area as the reference area for CD, CL, CM etc. by using the Plot-Reference-Area command in the top menu.



Figure-12, AeroWindTunnel Manual Data display selection.


Figure-13, AeroWindTunnel Manual Input mode used to define the geometry of a glider for flight analysis


Figure-14, AeroWindTunnel designed glider in flight.
 

4) MACH 1.4 JET AIRPLANE (Slider-Input) Back
To operate the example of a supersonic jet airplane operating at Mach 1.4 click the project file, FIGHTER PROJECT_FT.DAT using the main screen File menu. Then, when prompted click FIGHTER_AIRCRAFT_FT.TXT from the Fuselage Geometry and Wing-Fuselage Aerodynamic Center (ac) Location screen. Select Camouflage paint on aluminum in the Surface-Roughness pull-down menu. Close this screen when ready. To view coefficient plots click Plot-Coefficients on the main screen to display the Coefficient Plots screen. Select the reference area by clicking Total Wing Area in the Plot-Reference-Area command located in the pull-down menu of the Coefficient Plots screen. Then, click Plot-Experimental-Data in the top menu. In the File menu, click Load Experimental Data and then click FIGHTER_DATA.EXP, the wind tunnel experimental data from the book, Aircraft Design: A Conceptual Approach by D. P. Raymer. Finally, click PLOT and then CLOSE to see the plots.

FILES REQUIRED (FPS UNITS)
FIGHTER PROJECT_FT.DAT (Project file)
FIGHTER_AIRCRAFT_FT.TXT (Fuselage geometry file)
FIGHTER_DATA.EXP (Experimental data file)

FILES REQUIRED (MKS UNITS)
FIGHTER PROJECT_M.DAT (Project file)
FIGHTER_AIRCRAFT_M.TXT (Fuselage geometry file)
FIGHTER_DATA.EXP (Experimental data file)

Figure-15: Jet airplane zero-lift drag coefficient (CD0) compared with data from Aircraft Design: A Conceptual Approach by D. P. Raymer. Blue line represents AeroWindTunnel results for CD0 and red dot-line represents data from Aircraft Design



Figure-16,
AeroWindTunnel main screen used to define the basic  geometry of the supersonic jet airplane.



Figure-17,
AeroWindTunnel Fuselage geometry import screen. Simply define 20 plan-view and side-view station locations.


Figure-18, AeroWindTunnel results compared with zero-lift drag data from Aircraft Design for the supersonic jet airplane: A Conceptual Approach by D. P. Raymer. Blue line represents AeroWindTunnel results for CD0 and red dot-line represents data from Aircraft Design.

 

AeroCFD 2D-WING bonus feature addition to AeroWindTunnel Back

AeroWindTunnel includes a new program called 2D-WING for the aerodynamic analysis of two-dimensional (2-D) and finite aspect ratio (AR) airfoil sections. 2D-WING uses vortex lift panels to compute CD, CL and Cm,c/4 for airfoil sections using NACA four digit airfoils, streamlined, flat plate, double wedge (D'Wedge) and imported custom shapes for a wide range of 2-D and finite AR airfoils. Several NACA five-digit airfoils from Appendix III in the book Theory of Wing Sections allow the user to rapidly specify complex imported shapes. Other useful input variables include wing Reynolds number (Re) and angle of attack in degrees. Also, AeroCFD 2D-WING produces filled color contour plots and line color contour plots for pressure coefficient (Cp) and U/U0 where the number of contour levels can be specified from 3 to 256 levels. In addition, the following standard plots are produced, Cp verses chord length and U/U0 verses chord length for the upper and lower airfoil surfaces. Also, CL verses AOA, CD verses AOA, CD verses CL, CL/CD verses AOA and Cm verses AOA are quickly plotted. Finally, the total number of 2-D vortex panels that define the upper and lower surfaces of an airfoil can be specified as 100, 200 or 300.


Figure-19, A bonus feature of AeroWindTunnel is that AeroCFD 2D-WING is included at no extra cost to determine CD, CL and Cm for airfoils



Figure 20, Fin Vortex Panel Method Analysis Screen, Cp (pressure coefficient) Line Contours and Cp verses X.

AeroCFD 2D-WING uses 2D vortex lift panels to determine drag coefficient (CD), lift coefficient (CL) and moment coefficient (Cm,c/4) of airfoil sections. Airfoil section shapes are specified using the NACA four digit series, Streamlined, Flat Plate, D'Wedge and Imported shape options. The following steps outline the basic procedure used to operate AeroCFD 2D-WING.

Step 1.
The specification of Streamlined, Flat Plate and D'Wedge airfoil section shapes requires the specification of fin thickness in terms of maximum thickness in percent chord (Tmax/Chord X 100). Maximum fin thickness in percent chord for all airfoil types is specified by using the third data entry box on the first line, NACA Four-Digit series airfoil description. The first two data entry boxes for Streamlined, Flat Plate and D'Wedge airfoil section shapes are disabled. However, for the NACA Four-Digit series, the first two data entry boxes are enabled, where the first data entry box refers to maximum camber in percent chord and the second data entry box refers to position of the maximum camber in tenths of a chord from the leading edge (LE).

The following definitions are needed to define camber and camber location for NACA airfoils. First, the mean camber line is the locus of points halfway between the upper and lower surfaces of the airfoil as measured perpendicular to the mean camber line. Then, the chord line is a straight line that connects the leading and trailing edges of the airfoil and is simply referred to as the chord of the airfoil and is usually defined using the symbol, c. Using these definitions the camber is the maximum perpendicular distance between the mean camber line and the chord line of the airfoil. Camber location is simply located as a percentage of the chord length from the leading edge of the airfoil.

The complete specification of four-digit NACA airfoils and standard airfoils are summarized below for the first line, NACA Four-Digit series airfoil description. Where the first two spaces are disabled for Streamlined, Flat Plate, and D'Wedge airfoil section shapes but are required for NACA four-Digit airfoils.
[Max camber in percent chord], [Position of max camber in 1/10th chord], [Max thickness in percent chord].

Step 2. The Reynolds number (Re =
rUc/m) of the fin is defined on the second line of input data. Reynolds number is defined as the ratio of the inertial forces represented by the density of the medium (r), free stream velocity (U), and fin dimension (c) to the friction forces in the boundary layer represented by the viscosity of the medium (m). Reynolds number is automatically inserted if entering from AeroWindTunnel but is an input if entering from the AeroCFD 2D-WING command buttons. The following Reynolds number calculator is useful for computing Reynolds number for AeroCFD 2D-WING (Note: This off-site calculator has not been checked for accuracy). For more information and theory about Reynolds number please visit the Wikipedia on-line encyclopedia.

Step 3. Fin angle of attack (
a) is defined relative to the chord line for all section shapes on the third line of input data.

Step 4. Fin aspect ratio (AR = Span/Chord) is defined on the fourth line of input data. Aspect ratio must be non-zero and Checked to be included in the computation of CD, CL and Cm,c/4. The Aspect ratio input allows an "approximate solution" of end effects and 3D wings.

Step 5. Select one of five fin section shapes using the Airfoil Shapes pull-down menu. Streamlined, NACA four-digit, Flat Plate and D'Wedge section shapes are directly drawn after selection. In addition, arbitrary fin section shapes may be defined using the Import X-Y command. Many NACA Five-Digit section shapes are included in NACA_AIRFOILS.zip (located in the AeroWindTunnel directory) and are drawn by using the Import X-Y command after unzipping the file.

Import File format for each station (X) and ordinate (Y) given in percent of airfoil chord (LE is the Leading Edge): [Upper Surface X-location from LE], [Upper Surface Y-location from Chord Line], [Lower Surface X-location from LE], [Lower Surface Y-location from Chord Line] for each station from the LE to TE.

Step 6. Perform a 2D Vortex Panel aerodynamics analysis by clicking the SOLVE command button and follow the instructions displayed in the lower left status bar. Instructions displayed in the Status bar will state when a valid solution is achieved and when it is permissible to click the various plot command buttons.

Step 7. Display results using the following commands in the Plots pull-down menu: Cp verses X, U/U0 verses X, CL verses AOA, CD verses AOA, CD verses CL, CL/CD verses AOA, Cm verses AOA, U/U0 Contours (Filled and Line) and finally Cp Contours (Filled and Line). Where AOA refers to angle of attack in degrees, Cp = (P - PINF) / q = 1 - (U/U0)^2 is the pressure coefficient and U0 refers to the free stream velocity. Please note the results obtained by modifying data in AeroCFD 2D-WING are not reflected back into the Rocket analysis except for the number of panels that define the airfoil section. However, fin shapes used on the Fin and Launch Lug Geometry screen may be analyzed and the results plotted in AeroCFD 2D-WING. Reference: THEORY OF WING SECTIONS, by Abbott and Doenhoff.

NOTES:
1) Reference: THEORY OF WING SECTIONS, by Abbott and Doenhoff.
2) AeroCFD 2D-WING results (red dots) generated using the Save Results As command under File. The Results were plotted using Excel (or any spreadsheet program) and compared to THEORY OF WING SECTIONS data for the NACA 0012 and NACA 63-212 wing sections.

(1) NACA 0012 AIRFOIL VALIDATION
     
Figure 21, NACA 0012 CL verses AOA and CD verses CL.

(2) NACA 63-212 AIRFOIL VALIDATION
     

Figure 22, NACA 63-212 CL verses AOA and CD verses CL
 

REFERENCES
Introduction to Flight, John D. Anderson, 1989
Fundamentals of Aerodynamics, John D. Anderson, 1984
Computational Fluid Dynamics, The Basics with Applications, John D. Anderson, 1995
Modern Compressible Flow with Historical Perspective, John D. Anderson, 1982
Hypersonic and High Temperature Gas Dynamics, John D. Anderson, 1989
AIAA Design Engineers Guide, 4th Edition, 1998
Dynamics of Atmospheric Flight, Bernard Etkin, 1972
Aircraft Design: A Conceptual Approach, Daniel P. Raymer, 1989
Airplane Design, Part VI, Preliminary Calculation of Aerodynamic Thrust and Power Characteristics, Jan Roskam
Airplane Flight Dynamics and Automatic Flight Controls, Part 1, Jan Roskam, 1995
Airplane Flight Dynamics and Automatic Flight Controls, Part 2, Jan Roskam, 1995
Fluid Dynamic Drag, S.F. Hoerner, 1965
Fluid-Dynamic Lift, S.F. Hoerner, 1975
Full list of AeroRocket references

BACK TO TOP

DISCLAIMER: The developer does not guarantee the accuracy of AeroWindTunnel or the information in books provided as references. The reference books and computer program should not be used as an authoritative source for aircraft design data or methods. However, every effort has been made to make AeroWindTunnel as accurate as possible through extensive validation.

BACK TO TOP

ACCESS TO ONLINE INSTRUCTIONS (VERSIONS PRIOR TO 6.4.0.4): The online instructions are accessed by clicking File then Online Instructions from the main screen menu. Internet access is required to see the instructions and is located at the following URL: http://www.aerorocket.com/AeroWindTunnel/AWT_Instructions.shtml

ShowHTML.exe
displays the online instructions and is accessed by clicking File then Online Instructions as explained above. For Windows XP ShowHTML.exe is located in c:/Program Files/AeroWindTunnel. However, for Windows 7 and Windows VISTA ShowHTML.exe is located in c:/Program Files (x86)/AeroWindTunnel.


AeroWindTunnel Sample Files
Glider_Examples.zip located in the AeroWindTunnel directory, provides sample project files, fuselage geometry files and experimental data files for results comparison. Please decompress the examples file using WinZip as usual and locate all files in the AeroWindTunnel directory. For Windows XP the AeroWindTunnel directory is
located in c:/Program Files. However, for Windows 7 and Windows VISTA the AeroWindTunnel directory is located in c:/Program Files (x86).

AeroWindTunnel Sample Files
Glider_Examples.zip located in the AeroWindTunnel directory, provides sample project files, fuselage geometry files and experimental data files for results comparison. Please decompress the examples file using WinZip as usual and locate all files in the AeroWindTunnel directory. For Windows XP the AeroWindTunnel directory is
located in c:/Program Files. However, for Windows 7 and Windows VISTA the AeroWindTunnel directory is located in c:/Program Files (x86).

BACK TO TOP

 

SUMMARY OF FEATURES
1. Methods of data input and model definition
a) Aerodynamic components include fuselage, wings (standard and double-delta), horizontal tail, elevator and a user defined quantity of vertical fins.
b) Specify non-symmetric fuselage plan-view and side-view shapes using only 20 X-Y points arrayed in text (TXT) file format. AeroWindTunnel does not require DXF and IGES format geometry files for operation because AeroWindTunnel is not a CAD program.
Instead, AeroWindTunnel is a computer-based, conceptual-design wind tunnel program that uses slider-bar entry and imported fuselage shapes to quickly estimate stability of airplanes and gliders.
c) Results from 0.0 Mach to Mach 30 and angles of attack from -45 to +45 degrees
d) Manual entry of airplane flight coefficients from wind tunnel and CFD supplied data with complete display of values.
e) Slider-bar entry of airplane dimensions using the built-in ability to determine flight coefficients and flight derivatives with complete display of values.
f) Define atmospheric properties (pressure, density, viscosity etc) from Sea Level to 200,000 feet.
g) Velocity defined in meters/sec, feet/sec or Mach number.
h) For supersonic flow, wave drag includes propulsion inlet-area effects.
i) Geometry specifications include pointed nose, round nose, elliptical fuselage cross-section, rectangular fuselage cross-section, turbulent flow, laminar flow, and whether or not to include fuselage boat tail drag reduction.
j) Wave drag for wings having the following cross-sections may be specified: Single wedge (KLE=1), Symmetrical double wedge (KLE=4), Biconvex section (KLE=5.3), Streamline foil with x/c=50% (KLE=5.5), Round-nose foil with x/c=30% (KLE=6.0), Slender elliptical airfoil section (KLE=6.5) and finally Double wedge with maximum thickness at arbitrary x/c location (KLE=[c/x]/[1-x/c]).
k) Surface roughness effects for the following surfaces are used to determine the cut-off Reynolds number for each aerodynamic component: None (perfectly smooth), Camouflage paint on aluminum, smooth paint, production sheet metal, polished sheet metal, and finally smooth molded composites.
l) Aerodynamic effects include swept wing contribution to drag and lift.
m) Basic units of measurement may be specified as meters, centimeters, feet, or inches.
n) AeroWindTunnel instructions distributed in HTML format using WinZip compression.
o) Display WING or BODY drag divergence Mach number (MDD) depending on the effects of wing and body on wave drag.

2a. Input Flight coefficients and data
a) Flight altitude above sea level, Z
b) Main wing exposed surface area, Sref
c) Main wing mean aerodynamic chord, c
d) Horizontal tail exposed surface area, St
e) Vertical fin exposed surface area, Sf
f) Number of vertical fins, N_fins
g) Wing-body cg location from wing leading edge, h, c=ref
h) Distance from airplane cg to tail aerodynamic center, l_t
i) Airplane fineness ratio, L/D
j) Wing dihedral angle (+ up)
k) Main wing setting angle, -TE up, i_w
l) Tail dihedral angle (+ up)
m) Horizontal tail setting angle, +TE up, i_t
n) Airplane velocity, M/SEC, FT/SEC and Mach
o) Airplane mass, KG, LB
p) Geometric angle of attack, deg
q) Flight angle for unaccelerated flight, deg
r) Elevator deflection, +TE down, deg

2b. Input Flight coefficients and data
Wind Tunnel or CFD supplied data (All required if using Manual Input)
a) Wing-body aerodynamic center from wing LE, h_ac_wb, c=ref
b) Zero-lift angle of attack,
a @ L=0
c) Wing-body downwash angle,
e (a=0)
d) Wing-body down-wash gradient,
e_a
e) Wing-body moment coefficient, Cm
f) Wing-body lift-slope coefficient, CL
a
g) Horizontal tail lift-slope coefficient (if required), CL
a
h) Elevator effectiveness lift-slope, CL
de
i) Tailless elevator moment (if required), Cm
de

3. Output Flight coefficients and parameters
a) Slope of pitch-moment coefficient, Cm
a
b) Pitch-moment coefficient around cg at initial angle of attack, Cm
c) Slope of yaw-moment coefficient due to sideslip, Cn
b
d) Slope of rolling-moment coefficient due to sideslip, Cl
b
e) Absolute angle of attack for trimmed flight (Cm = 0),
aa
f) Moment coefficient around cg when lift (L) = 0, Cm, 0
g) Tail volume ratio, VH
h) Neutral point location for pitch measured from wing leading edge (LE) and normalized by main wing mean aerodynamic chord (c), hn.
i) Airplane static margin for pitch: SM = hn - h where h is the wing-body cg location from the wing LE.
j) Airplane neutral point location for yaw measured from the nose-tip and normalized by the vertical fin mean aerodynamic chord (cf), hn_fb.
k) Airplane static margin for yaw, SM = hn_fb - h where h is the wing-body cg location from the nose-tip.
l) Wing lift slope for pitch, CL
a_w
m) Body lift slope for pitch, CL
a_body
n) Horizontal tail lift slope for pitch, CL
a_tail
o) Vertical fin lift slope for yaw, CL
a_fin
p) Rate of change of lift coefficient with elevator deflection, CLd
q) Rate of change of moment coefficient with elevator deflection, Cm
d

4. Output requirements for climbing and level flight
a) Airplane drag coefficient, CD
b) Airplane lift coefficient, CL
c) Airplane lift to drag ratio, L/D
d) CL required for level flight
e) Thrust required for climbing and level flight
f) Angle of attack for level unaccelerated flight (deg)
g) Elevator angle at trim, +TE down (deg)
h) Climb Rate in meters/min or feet/min for Flight angle

5. File and data manipulation
a) Plot total airplane moment coefficient (Cmcg) around center of gravity verses angle of attack (
a).
b) Plot wing-body moment coefficient (Cmcg_wb) around center of gravity verses angle of attack (
a).
c) Save project files as .DAT files to disk.

d) Save results to disk as .DAT files.
e) Print screen images to the printer.

6. Real-time airplane plan view and side view with the following information
a) Center of gravity (Cg) location marked on plan/side view as a
green dot.
b) Wing-body aerodynamic center (ac_wb) location marked on plan view as a
red dot.
c) Total airplane aerodynamic center (ac_pitch) or neutral point marked on plan view as a black dot.
d) Tail aerodynamic center (ac_tail) location marked on plan view as a red dot.
e) Vertical-fin-body aerodynamic center (ac_yaw) location marked on side-view as a black dot.
f)  Vertical-fin aerodynamic center (ac_fin) location marked on side-view as a red dot.

7. Real-time airplane design values displayed on plan view and side view
a) Airplane center of gravity (cg).
b) Airplane pitch aerodynamic center or neutral point location (ac_pitch).
c) Wing-body aerodynamic center location (ac_wb).
d) Exposed-wing aspect ratio (AR_wing).
e) Tail aerodynamic center location (ac_tail).
f) Airplane yaw aerodynamic center or neutral point location (ac_yaw).
g) Fin aerodynamic center location (ac_fin).
h) Horizontal Tail aspect ratio (AR_tail).
i) Vertical Fin aspect ratio (AR_fin).

8. Results Plots
a) Plot nine coefficients verses angle of attack (-25 to 25 degrees).
b) Plot nine coefficients verses Mach number (0.0 to Mach 30).
c) Plot coefficients using exposed wing area, total wing area, body planform area or maximum body frontal area as plot reference.
d) Insert experimental data into coefficient plots for comparison.
e) Recent enhancements include the ability to display
3-D Orthographic images of aircraft model geometry.

9. Much more ...


AEROWINDTUNNEL REVISIONS
AeroWindTunnel 6.4.0.5 (2/10/12)

1) Increased maximum Mach number from Mach 4 to Mach 30 for all coefficient plots verses Mach number and angle of attack.
2) Improved airframe hypersonic flow (greater than Mach 5) drag (Cd) and lift (CL) correlation with experimental aerodynamic data.
a) Improved Cd verses Mach number comparison with text book aerodynamic data for the Fighter Airplane Project.
b) Improved Cd verses Mach number comparison with test results from the AeroRocket supersonic wind tunnel and VisualCFD results for the HTV-3X project.
c) Improved supersonic analysis prediction for airframe lift slope, CN
a and center of pressure location, Xcp for all supersonic Projects.
3) Fixed Slider-input "ON" and Import-Fuselage "OFF" error where Airplane velocity (V2) increases orders of magnitude larger than the saved value for V2 when opening a Project.
4) Added the ability to save a file with the STK extension for importing Cd verses Mach number information into the Satellite Trajectory Kit routine recently added to StarTravel.
5) Added the ability to compute exact Maximum Body Frontal Area (Smax) by specifying a Section-Factor area correction for the elliptical and rectangular cross-sections used by AeroWindTunnel.
6) In the RESULTS section on the Fuselage Geometry and Wing-Fuselage Aerodynamics Center Location screen, changed the text, Drag coefficient at aoa=0:CD_0 =, to Skin friction drag coefficient: CD_0 =, for added clarity.
7) In the RESULTS section on the same screen declared that center of gravity location (h_cg), vertical fin aerodynamic center location (h_ac_fin), pitch aerodynamic center location (h_ac_pitch), yaw aerodynamic center location (h_ac_yaw), horizontal tail aerodynamic center location (h_ac_tail), Wing-body aerodynamic center location (h_ac_wb), Wing-body aerodynamic center location (h_ac_wb) and Fuselage pitch/yaw aerodynamic center locations (h_ac_b) are normalized by airframe length. In addition, the tool-tip also declares that Skin friction drag (CD_0), Base drag (CD_base) and Wave drag (CD_wave) values are referenced to the exposed wing area.
8) Due to the extent of these modifications and error fixes users who purchased AeroWindTunnel starting with version 6.4.0.1 can receive a free upgrade to the newest version. In order to verify the validity of each request for a FREE version of AeroWindTunnel please provide name, the exact email used during original purchase and date of original purchase.

AeroWindTunnel 6.4.0.4 (3/07/10)

1) For AeroWindTunnel 6.4.0.4
the online instructions are accessed by clicking File then Online Instructions then selecting either When Using Windows XP or When Using Windows 7 or VISTA. Previously when operating under Windows 7 and VISTA the message, Error displaying Online Instructions was displayed when trying to access the online instructions. This occurred because under Windows XP, ShowHTML.exe is located in c:/Program Files/AeroWindTunnel but for Windows 7 and Windows VISTA the routine is located in c:/Program Files (x86)/AeroWindTunnel.

AeroWindTunnel 6.4.0.3 (2/07/10)

1) Previously, in the Coefficient Plots Verses Angle of Attack screen when the Save Coefficient Data As  command is selected from the File pull-down menu the values for Thrust Required (TR) and Velocity Required (VR) for level flight displayed in the Cd verses angle of attack listing were smaller than actual by 10^-6. Now, the actual value is the displayed value multiplied by 1000 as described in the heading for TR and VR.

2) Previously, in the Plots pull-down menu when the Coefficients vs. Mach number command is selected the listing did not display coefficients as a function of Mach number but incorrectly showed the variation as a function of angle of attack in degrees.

AeroWindTunnel 6.4.0.1/6.4.0.2 (4/22/09)

1) Added the ability to determine Thrust Required for Climbing Flight and level flight. In previous versions only thrust required for level flight was computed. In addition, made Flight angle an input variable where Climb Rate is displayed in feet/minute or meters/minute depending on initial units. Validated these modifications using the existing collection of airplane analyses in addition to a new analysis for the Me-163 rocket plane developed during World War II by Germany. AeroWindTunnel results for drag coefficient (CD) and climb speed compare exactly with results found in Fluid Dynamic Drag by S.F. Hoerner on pages 14-9 and 14-10 where in TABLE-A, CD = 0.012  and in Figure-7, Climb Rate = 11,600 feet per minute for the revolutionary rocket plane. Please request the new Me-163 analysis if you already own AeroWindTunnel.

2) Corrected a CGS and IPS units error. MKS and FPS units were not affected. This units error effected supersonic flight coefficients and thrust requirements.

AeroWindTunnel 6.3.0.0 (5/24/08)

1) Added double-delta wing geometry to the Fuselage Geometry and Wing-Fuselage Aerodynamic-Center screen. After the double delta wing is specified the user has the option of inserting a double-delta wing into the plots or inserting an equivalent double-delta wing into the analysis and plots. The double-delta wing and its variants are used to reduce the affect of the rearward aerodynamic-center shift that occurs in the transition between subsonic and supersonic flight.

AeroWindTunnel 6.2.0.2 (4/06/08)

1) Vertical fins could not be located at wing-tips and there was no fin vertical control. Now, fin-to-fin spacing and fin Y-location inputs can locate one vertical fin at each wing-tip location. This version adds a vertical fin Y-location input to accurately locate up to two fins in the vertical direction.
2) Wing span appeared to change for imported models with variable fuselage diameter as the wing location was modified using the LE location from nose tip slider-bar control. This was a geometry problem that did not alter AeroWindTunnel results.
3) Effects for some data inputs on the Fuselage Geometry screen were not seen immediately in the 3-D wire frame plots.
4) Added the HTV-3X or BlackSwift, XCOR's Lynx and AeroRocket's AeroEagle to AeroWindTunnel's collection of project files in Glider_Examples.zip. Already included in the AeroWindTunnel collection are project files for the X-30 NASP, HL-20 and F-16 type jet airplane in addition to several other gliders and examples from text books.

AeroWindTunnel 6.2.0.1 (12/17/07)
1) For the Plot Coefficients screen, Cd verses angle of attack and Cd verses Mach number are now displayed from 0.0 to a maximum value.
2) For the Plot Coefficients screen, plot error occurred for all coefficients verses Mach number if geometric angle of attack equaled zero-lift angle of attack.
3) Run-time error "6" overflow occurred for manual input if most values were made zero in the MANUAL ENTRY: WIND TUNNEL DATA section.
4) Input blocks for Horizontal tail exposed surface area and Vertical fin exposed surface area were sometimes zeroed when manual input was specified.

SYSTEM REQUIREMENTS
(1) Screen resolution: 1024 X 768
(2) System: Windows 98, 2000, XP, Vista, Windows 7, NT or Mac with emulation
(3) Processor Speed: Pentium 3 or 4
(4) Memory: 64 MB RAM
(5) English (United States) Language

Please note this web page requires your browser to have
Symbol fonts to properly display Greek letters (
a, m, p, and w)

BACK TO TOP

| MAIN PAGE | PRODUCTS | CONSULTING | MISSION | RESUME | ROCKETRY |