mae 4350 - lab 4 aerodynamics

MAE 4350 – Aerospace Vehicle Design I

Fall Semester 2009

Lab 4 Aerodynamic Estimation

4.1 Function of Aerodynamics in Design4 2 Aerodynamic tool box introduction4.2 Aerodynamic tool box introduction4.3 Aerodynamics for Performance

Methodology4.4 Aerodynamics for Stability and Control

Methodology4 5 A d i f St t4.5 Aerodynamics for Structures

Methodology4.6 Assignment

Gary Coleman

AVD Laboratory

November 2009 © G. Coleman / UTA MAE / AVD Lab

November 6, 2009

Mechanical and Aerospace Engineering Department (MAE)

3 Initial Geometry, Weight and Balance

Aerodynamic Estimation

Geometry, Weight and Balance Is typically derived duringthe configuration layout phase for the basic trades-studies oninterest. The methods employed are typically statistical innature and serve only as a start point for the design process

Covered In Lecture

Responsible Teams: CAD and Synthesis (Chief Engineers)

Derivation of initial Geometry Weight &

Balance

Market

Mission Definition

Parametric Sizing

ConfigurationLayout

Balance

Covered In Lecture

Configuration Evaluation

Flight Simulation

ProductReview

Lab Section: Disciplinary Methods Introduction


4 Aerodynamic Estimation for Performance


Configuration Trade-studies

Initial weight & balance and geometry

Configuration Evaluations Process

Aerodynamic estimation

Propulsion estimation

Stability and control analysis

g g y

Structural analysis

Performance analysis

Internal systems analysis

Cost analysis

Convergence CheckExample: Compare initial and final values for weight

Revised weight & balance estimation

Example: Compare initial and final values for weight

Present Trade-study results


4.1 Function of Aerodynamics in Design


Aerodynamics: Is the prediction and tailoring of theaerodynamic forces and moments required for predictingthe aircrafts,

1. Performance2. Stability and Control3. Structural Loads

Performance:

R i th d l M i lift ffi i t f hRequires the drag polar, Maximum lift coefficient for each mission segment and Lift curve


Stability and Control:


Requires static, dynamic and control derivatives for critical flight conditions

Configuration Derivatives

Ref: NASA CR-2144


Ref: NASA CR 2144

Structure:

R i d i l d di t ib ti h


Requires aerodynamic load distribution over eachconfiguration component for the critical load cases

Pressure Distribution

Ref: Howe, “Aircraft Loading and Structural Layout,” 2004


4.3 Example: Citation X


Flight Conditions and Configuration settings:

For this example the Take-off flight condition will be examined. From the Mission Specification and Parametric sizing results the flight conditions and configuration settings areflight conditions and configuration settings are.

Altitude: Sea-levelVelocity: 137 kts (231 ft/s)Mach: 0.15Re: 1.45x106

flap: 15.0Landing Gear: Down


4.2 Aerodynamic tool box introduction


2-D aerodynamics:Airfoil Drag Polar and Lift Curve:

1) Look for experimental data1) Theory of Wing Sections2) UIUC Ai f il D t Sit2) UIUC Airfoil Data Site

(http://www.ae.uiuc.edu/m-selig/ads.html )3) Paper DATCOM4) Google!

2) Numerically Prediction• Digital DATCOM – Method of singularities,

corrected for viscous and compressibility effects• EPPLER – Potential flow solver• X-FOIL– Potential flow solver corrected forX-FOIL– Potential flow solver, corrected for

compressibility • JavaFoil – Potential flow solver, corrected for

compressibility • TSFOIL – Transonic small disturbance theory

Fl t C i l CFD ft ( t il bl• Fluent – Commercial CFD software (not available in Capstone Lab)




3-D aerodynamics:Drag Polar, Lift Curve, aerodynamic loads and stability and

control derviatves:1) Hand-book Methods

2) Digital DATCOM

1) AVL Vortex Lattice Code




Flight Conditions and Configuration settings for toolbox description:

For this example the Take-off flight condition will be examined. From the Mission Specification and Parametric sizing results theFrom the Mission Specification and Parametric sizing results the flight conditions and configuration settings are.

Altitude: Sea-levelVelocity: 137 kts (231 ft/s)Mach: 0 15Mach: 0.15Re: 1.45x106

flap: 15.0Landing Gear: Down


2-D Aerodynamics:


Citation X Wing Airfoil Approximation:The Citation X’s wing is composed of supercritical airfoils

which vary from root to tip. However, the actual airfoils are not availbile in the public domain and therefore the airfoils can beavailbile in the public domain and therefore the airfoils can be approximated as follows

-0.1

-0.05

0

0.05

0.1

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1

x/c

y/c

Upper

Lower

Mean camber line

GIII BL 145

GIII BL 45 9750 mm.

Cessna 7500

-0.1

-0.05

0

0.05

0.1

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1

x/c

y/c

Upper

Lower

Mean camber line

GIII BL 45

F th f thi L b th ti i i i t d

-0.1000

-0.0500

0.0000

0.0500

0.1000

0.00 0.10 0.20 0.30 0.40 0.50 0.60 0.70 0.80 0.90 1.00

x/c

y/c

Upper

Lower

Mean camber line

4900 mm.

For the purposes of this Lab the entire wing is approximated with the 10% t/c GIII BL 45 from the Gulfstream III. The airfoil ordinates can be found from the UIUC Airfoil Data Site (http://www.ae.uiuc.edu/m-selig/ads.html )


2-D Aerodynamics:


Citation X Empennage Airfoil Approximation:The Citation X’s empennage also composed of supercritical

airfoils with an approximate t/c of 10 % for the vertical and 8% for the horizontal.

For the purposes of this lab the NACA 64a010 and NACA 64-008a are used. The ordinates can be found from the UIUC Airfoil Data Site (http://www.ae.uiuc.edu/m-selig/ads.html )

0.1Upper

Lower

Mean camber line

NACA 64a010

-0.1

-0.05

0

0.05

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1

x/c

y/c

Mean camber line

NACA 64 008a

0 1

-0.05

0

0.05

0.1

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1

y/c

Upper

Lower

Mean camber line

NACA 64-008a

-0.1

x/c


2-D Aerodynamics:


Airfoil Drag Polar and Lift Curve:1) Look for experimental data

1) Theory of Wing Sections2) UIUC Ai f il D t Sit2) UIUC Airfoil Data Site

(http://www.ae.uiuc.edu/m-selig/ads.html )3) Paper DATCOM4) Google!

2) Numerically Prediction• Digital DATCOM – Method of singularities,

corrected for viscous and compressibility effects• EPPLER – Potential flow solver• X-FOIL– Potential flow solver corrected forX-FOIL– Potential flow solver, corrected for

compressibility • JavaFoil – Potential flow solver, corrected for

compressibility • TSFOIL – Transonic small disturbance theory

Fl t C i l CFD ft ( t il bl• Fluent – Commercial CFD software (not available in Capstone Lab)

These will be used for the Capstone Lab


2-D Aerodynamics: X-FOIL


X-FOIL is a menu based Airfoil analysis and design program developed by Mark Drela at MIT and is available for free under a GNU General Public License.

Web site: http://web.mit.edu/drela/Public/web/xfoil/

Wing airfoil

1) Set-up airfoil coordinate file X-Foil\GIIIBL45 dat1) Set-up airfoil coordinate file X-Foil\GIIIBL45.dat

Notes: • first line is a the airfoil name• Use only spaces between columns

C di t t t t th t ili d d• Coordinates start at the trailing edge and run forward only the top surface and aft along the bottom surface

0.05

0.1Upper

Lower

Mean camber line

-0.1

-0.05

0

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1

x/c

y/c




2) Place the coordinate data file into the same folder as the X-FOIL executables. Double click xfoil.exe. You will see a directory and command listing.

3) T pe ‘LOAD GIIIBL45 dat’3) Type ‘LOAD GIIIBL45.dat’


2-D Aerodynamics: X-FOIL4) Change to the Geometry design routine by typing


4) Change to the Geometry design routine by typing ‘GDES’ and hitting return. A plot visualizing the airfoil will appear

In this case X-FOIL gave a warning message that the airfoil has a poor coordinate distribution. To correct thi th CADD d d t d th l lthis the CADD command was used to reduce the local panel angles.

5) Type ‘CADD’ in the GDES directory. This this case the function was used twice to reduce the maximum panel angle was around 3 deg.

6) Hit the ‘Enter’ to return to the XFOIL directory

7) Type ‘PCOP’ To re-panel the airfoil according to the


7) Type PCOP To re panel the airfoil according to the new coordinate points. This refinement will increase the accuracy of the final pressure distribution and drag polar



8) Change to the Operation routine by typing ‘OPER’ from the X-FOIL directory. (Type ‘?’ to show the directory and command list again)

9) Turn on the Viscous mode and input the Reynolds and9) Turn on the Viscous mode and input the Reynolds and Mach number during take-off.1) Type VISC and following the prompts2) Type MACH and follow the prompts

10) T th t i t l ti f ti Thi ill10) Turn on the auto point accumulation function. This will produce an output file with the drag polar results.1) Type ‘PACC’

1) Provide a name for the drag polar file2) Provide a name for the output dump file

11) Specify range of angle of attack1) Type ‘ASEQ’ and follow the prompts.




12) Check the solution1) Check for sharp spikes (mostly due to numerical

instabilities!!)

2) 1st try increasing the viscous iteration limit with2) 1 try increasing the viscous iteration limit with the ‘ITER’ command. If the problem presists try refining the number of panels and panel angles.

Note: Supercritical airfoils are typically difficult to model with panel methods!to model with panel methods!




13) Plot drag polar1) Double click ‘pplot.exe’ in the X-Foil folder2) Type ‘1’ the read the drag polar file and enter the

name of the drag polar file created earlier. Hit return twicereturn twice

3) Type ‘3’ to plot the drag polar

14) The drag polar data file can also be cop and pasted in14) The drag polar data file can also be copy and pasted in Excel for further formatting.

15) Repeat for the empennage airfoils


3-D Aerodynamics: Hand calculationsBefore running any computer code it is important to have a


Before running any computer code it is important to have a sanity check. The simplified handbook collected in methods described in Approximate Drag Polar Method.doc can be used for a quick approximation of the drag polar and maximum lift coefficient.

Notes• These methods are approximate in nature and

should only be used for parametric sizing purposes or for a sanity checkpu poses o o a sa ty c ec

• Most of these methods come from the USAF DATCOM, AIAA Aerospace designer engineers guide and Roskam Part I.


3-D Aerodynamics: Digital DATCOMDigital DATCOM is Digital version of the USAF DATCOM


Digital DATCOM is Digital version of the USAF DATCOM semi-empirical hand-book methods for aerodynamic estimation. Digital DATCOM uses a simple text file input and output interface and has been a stable of aerodynamic prediction in conceptual design sense the late 1970.

Digital DATCOM is an open source FORTRAN 77 program and is available in the public domain

See the AVD Digital DATCOM Quick Tour and Digital See t e g ta CO Qu c ou a d g taDATCOM users manual to get started

AVD Digital DATCOM Quick Tour.docDigital DATCOM MANUAL.PDF

NotesNotes• The airfoil sections data can be input from the X-

Foil results or the airfoil coordinates can be input manually. In the later case Digital DATCOM uses a small disturbance method to predict the airfoil characteristics.

• Familiarize your self with the applicability of Digital DATCOM. Sense the tool is based on semi-empirical methods it cannot be applied to all configurations.g


3-D Aerodynamics: AVL Vortex Lattice Code


What is a vortex lattice code?

2-D case – approximating an airfoil as a flat plate at some angle of attack with a vortex filament, with strength , at the ¼ chord position and a control point at the ¾ chord position pos t o a d a co t o po t at t e ¾ c o d pos t o(Weissinger’s approximation).

The lift (L’) and down wash (wi) and corresponding down wash velocity can be computed using the “Biot-Savart Law” (See McCormick(5) or Dreier(6) for more detail)

To produce a chord wise lift distribution simple add more vortex filaments and control pointsfilaments and control points

[5] McCormick, B.W., “Aerodynamics, Aeronautics and Flight Mechanics,” Wiley, NewYork, 1979


York, 1979

[6] Dreier, M.E., “Introduction to Helicopter and Tiltrotor Flight Simulation,” AIAAEducational Series, American Institute of Aeronautics and Astronautics.,Reston, VA., 2007




3-D case – Expanding on the infinite flat plat assumption to a finite thin wing can be derived assuming two wing tip vortices extending from the wing quarter chord of each aft connected e te d g o t e g qua te c o d o eac a t co ectedwith a bounding vortex along the wing ¼ chord. Resulting in “Horse Shoe” vortex with strength .

Through summing the wash effects of each vortex at a central control point at mid span and ¾ chord the lift and induced dragcontrol point at mid span and ¾ chord the lift and induced drag for this wing can be approximated for small angles of attack. Moving this lift vector to the ¼ chord produces the wing pitching moment.





To produce a span wise lift distribution add more horse shoe vortices at the wing ¼ chord

To produce a chord wise and span wise distribution add horse shoe vortices at various chord locations




How to operate AVL: AVL is a command driven (similar to x-foil) which reads geometry and weight data from data files.

Getting Started: place “avl.exe” in the “runs” folder and follow the. Double click avl.exe t e oub e c c a e e

The remainder of this introduction is a visualization of the “session1.txt” file which summarizes the basic commands to operate AVLoperate AVL.




To Load and visualize the aircraft geometry:1) Load the geometry file “vanilla.avl”2) Change to the “.OPER” directory




To Load and visualize the aircraft geometry:3) Type “G” to bring up a wire frame plot of the geometry 4) Type “K” to bring up keyboard commands for

manipulating the plot




To Run AVL for a specified flight condition:1) From the “OPER” directory type “M” to modify the

flight conditions2) Enter the first letter of the variable you wish to specify

1) Example: “MN” for Mach number MN) a p e o ac u be


3) Repeat for velocity, density, Mass, gravitational constant, and center of gravity (minimum).



To Run AVL for a specified flight condition:4) AVL does not appear to run angle of attack or mach

number sweeps as done with DATCOM or Linair (If you find a way let me know!). There for you can specifically define the angle of attach, side-slip angle, etc., through de e t e a g e o attac , s de s p a g e, etc , t ougthe constraint table which appears in when you are in the “OPER” directory

5) Set the angle of attack to 4.0 degrees by typing “A” to select angle of attack. Then Select the variable which AVL will use to set the angle of attack. Select angle of attach by entering “A” and finally enter the angle of y g y gattack 4.0 deg

Note: AOA can be constrained by any variable listed (Like CL)




To Run AVL for a specified flight condition:6) Type “x” to execute the vortex lattice code. The

following screen will appear

6) Repeat the process for every flight condition or constraint




To Visualize Output with AVL:1) To Visualize the Lift distribution for all lifting surfaces

type “T” for the Trefftx plane plot.

Wing Lift Distribution

Horizontal Tail Lift Distribution

Downwash angle distribution




To Visualize Output with AVL: 1) To Visualize the pressure distribution for all lifting

surfaces type “G” to return to the geometry plot

1) Type “LO” to visualize the wing loading (pressure ) ype O to sua e t e g oad g (p essu edistribution)

Wing Lift Distribution






To Dump Output AVL to data files: 1) To output

1) Stability derivatives type “ST” or “SB”2) Total forces “FT”) ota o ces3) Surface forces “FN”

(wing, horizontal tail, etc.)4) Strip forces (lift distribution) “FS”5) Element forces (control points) “FE”6) Strip shear moments “VM”6) Strip shear, moments VM

(drag and pitching moments) 7) Hinge moments “HM”

2) For each command it will prompt you to provide a file name to write the output.






To Build an AVL model of the Citation X: 1) Start with the “bd2.avl” as a template for a wing, body

horizontal and vertical tail configuration.

2) Modify the reference areas mach number center of ) od y t e e e e ce a eas ac u be ce te ogravity references and CD0 (Vortex lattice methods only predict induced drag!!!!!!)

3) Modify the fuselage according the description in AVL’s Users guide (avl doc txt)Users guide (avl_doc.txt)Notes:1) Fuselages and nacelles are modeled as

uncambered bodies of revolution with circular cross-sections. Therefore, for the Citation X use the Top view to determine the cross-sectional radius at each x-station.

2) Fuselage bodies are input in the same manner airfoils (i.e. specify radii from tail to nose across the top of the fuselage followed by the radii from p g ynose to tail along the bottom surface

4) For each lifting surface (wing, horizontal tail, vertical tail, etc. specify

5) For each command it will prompt you to provide a file name to write the output.




To Build an AVL model of the Citation X: 4) For each lifting surface (wing, horizontal tail, vertical

tail, etc.) any number of chordwise locations (root, tip, mid-span, etc.) specify Leading edge x, y, z location, chord length, incidence angle and airfoil ordinate file c o d e gt , c de ce a g e a d a o o d ate eaccording to the AVL users guide.

Notes:1) The AVL can read the same airfoil data files as x-

foil2) You may use one or more airfoils for this model2) You may use one or more airfoils for this model.

-0.1

-0.05

0

0.05

0.1

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1

x/c

y/c

Upper

Lower

Mean camber line

GIII BL 145

XLE3

yw

-0.1

-0.05

0

0.05

0.1

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1

x/c

y/c

Upper

Lower

Mean camber line

GIII BL 45

YLE3

X

-0.1000

-0.0500

0.0000

0.0500

0.1000

0.00 0.10 0.20 0.30 0.40 0.50 0.60 0.70 0.80 0.90 1.00

x/c

y/c

Upper

Lower

Mean camber line

Cessna 7500

Yw

YLE2

XLE2

xw

Xw

zw

i3


Zw

iwxw

i3

ZLE3

4.3 Aerodynamics for performance methodology


Performance Team Aerodynamic Requirements:The performance team requires the aerodynamic dragpolar and lift curve to predict range, take-off and landingfield length, climb gradients, time to climb, service ceiling,etc.etc.

Minimum Deliverables:

• Drag Polar for each mission segment• Lift curve for each mission segment

Basic Procedure:

1. 2-D Airfoil selection/analysis (wind-tunnel data or X-FOIL)

2 3 D d b ild (H d b k DATCOM AVL)2. 3-D drag build-up (Hand-book, DATCOM, AVL)3. 3-D Lift curve with and without flaps (Hand-book,

DATCOM, AVL)

Recommended References:[1] Roskam, J, “Airplane Design, Part I: Preliminary Sizing of Airplanes,”

DARcorporation, Lawrence, Kansas, 2004

[2] Roskam, J, “Airplane Design, Part VI: Preliminary Calculation of Aerodynamic, Thrust, and Power Characteristics,” DARcorporation, Lawrence, Kansas, 2004

[3] Hoak, D.E, Finck, R.D., “USAF Stability and Control DATCOM,” Flight Control[3] Hoak, D.E, Finck, R.D., USAF Stability and Control DATCOM, Flight ControlDivision Airforce Flight Dynamics Laboratory, Wright-Patterson Air ForceBase, Ohio, 1978

[4] Hoerner, S.F,, “Fluid Dynamic Drag,” Midland Park, NJ, 1985

[5] Hoerner, S.F,, “Fluid-Dynamic Lift,” Midland Park, NJ, 1965


Performance Mission Segments:

The aerodynamics team must compute the trimmed lift and drag


The aerodynamics team must compute the trimmed lift and drag for the primary mission segments with the appropriate configuration settings

Mission requirements

Payload Weight (Kg)Payload Weight (Kg)

Crew (2) 184 kg (410 lbs)

Maximum Passengers (12) 1,110 kg (2,460 lbs)

Design Passengers (6) 600 kg (1320 lbs)

Range

Design (0.82 M) 5,740 km (3,100 nm)

High-Speed (0.92M) 4130 km (2,300 nm)

Velocity

High-speed cruise (mid-weight) 0.92 M

Design Cruise Speed 0.82 M

Altitude (m)

Max operating 15,000 m (49,000 ft)

Design Cruise (0.82 M) 15,000 m (49,000 ft)

Max cruise speed (0.92 M, mid-weight) 11,300 m (37,000 ft)

Take-off field length (TOGW) 1 570 m (5 140 ft)Take off field length (TOGW) 1,570 m (5,140 ft)

Landing field length (max landing weight) 1036 m (3,400 ft)

Fuel reserves 45 min at 1,524 km (5,000 ft)


4.4 Aerodynamics for Stability and Control methodology


S&C Team Aerodynamic Requirements:The S&C team must assess the stability and controllabilityof the aircraft and flight conditions which typically themost demining. These flight conditions are termed DesignConstraining Flight Conditions (DCFC)Constraining Flight Conditions (DCFC)


• Static stability derivatives at each DCFC• Dynamic stability derivatives at each DCFC

Control derivatives at each DCFC• Control derivatives at each DCFC

Basic Procedure:

1. Outline DCFC’s2 C t t i d t ti t bilit d i ti (DATOM2. Compute trimmed static stability derivatives (DATOM

/ AVL)3. Compute dynamic stability derivatives (DATOM / AVL)4. Compute control derivatives (DATOM / AVL)5. Produce look-up tables/figures for each DCFC

Recommended References:

[1] Roskam, J, “Airplane Design, Part VI: Preliminary Calculation of Aerodynamic, Thrust, and Power Characteristics,” DARcorporation, Lawrence, Kansas, 2004

[2] Hoak, D.E, Finck, R.D., “USAF Stability and Control DATCOM,” Flight ControlDivision Airforce Flight Dynamics Laboratory, Wright-Patterson Air ForceBase, Ohio, 1978

[3] Etkin B., Reid, L.D., “Dynamics of Flight: Stability and Control,” 3rd Edition, JohnWiley and Sons, Inc, New York, 1996

[3] Torenbeek E “Synthesis of Subsonic Airplane Design ” Delft University Press


[3] Torenbeek, E., “Synthesis of Subsonic Airplane Design,” Delft University Press,London, 1996

Design Constraining Flight Conditions (DCFC):

In general the following table can be used to define the proper


In general the following table can be used to define the proper DCFC’s for each control effector;

LoCE – Longitudinal Control Effector (Elevator)DiCE – Directional Control Effector (Rudder)LaCE – Lateral Control Effector (Aileron)( )

List of Classical DCFC’sDCFC Description

LoCELoCE

Trimmed Cruise Estimation of tim drag from the LoCE.

High ‘g’ Maneuvering

LoCE's ability to perform pull-up/push-over maneuvers at maximum g loading.

Take-off Rotation LoCE's ability to lift the nose of the ground at rotation speed.

High Low speed LoCE's ability to maintain trim at forward c g during low-speed landing approachHigh , Low speed LoCE s ability to maintain trim at forward c.g. during low speed landing approach with flaps-down, engines at idle, and high angle of attack.

DiCE

Crosswind Landing DiCE's ability to maintain straight ground path during take-off and landing

Anti-symmetric Power

DiCE's ability to maintain straight flight path with most outboard engine inoperable

Crosswind Landing with OEI

Combination of Cross-wind landing and Anti-symmetric power

Adverse Yaw DiCE's ability to compensate for yawing moments produced by the aileron during rolls or high a, low speed, steep coordinated turns.

LaCE

Roll Performance LaCE's ability to bank the aircraft to a required bank angle in the required timeRoll Performance LaCE s ability to bank the aircraft to a required bank angle in the required time.


Stability and Control Coefficients and Derivatives:


What does the S&C team need?- trimmed Stability and control derivatives

Longitudinal Lateral/Directional

Coefficients

Static Derivatives

0DC

DC LC

0LC0mC

mClC nC yC

trimDCtrimLC

trimmC

Static Derivatives

Dynamic Derivatives

Control Derivatives

DC LC mC

LoCEDC LoCELC LoCEmC

lC nC yC

uLCuDC

umCqLC qmC

LC mC plC

pnC pyCrl

C rnCryC

LaCElC LaCEnC LaCEyC

DiCElC DiCEnC DiCEyC

How to compute these parameters?Handbook Component Build-up Methods (Etkin, USAF DATCOM, Digital DATCOM)

Numerically Vortex Lattice Methods (AVL)

How to trim the configuration at each flight condition?Roskam Part VII: Roskam Trim.pdf

Trim is defined asLift = Weight

Thrust = DragPitching moment = 0

V

Lw

Lt

V

V’

D

Dt

M

Mact

lt


DwMacw

4.5 Aerodynamics Load Estimation


S&C Team Aerodynamic Requirements:Structures team must determine the structural concept,layout and weight for the aircraft. To accomplish theyrequire aerodynamic loads to estimate the forces andmoment on the structure for critical load casesmoment on the structure for critical load cases


• Distributed Aerodynamic Loads for the wing, fuselageand empennage for each critical load case (PressureDistribution)Distribution)

• Aerodynamic forces and moments (CL, CD, CM) foreach load case

Basic Procedure:

1. Outline critical load cases2. Compute disturbed Lift, Drag and Pitching Moment

for each aircraft component3. Produce look-up tables/figures for each critial load

casecase

Recommended References:

[1] Niu, M., “Airframe Structural Design,” Technical Book Company, California, 1990

[2] H S F “Fl id D i Lift ” Midl d P k NJ 1965[2] Hoerner, S.F,, “Fluid-Dynamic Lift,” Midland Park, NJ, 1965

[3] Roskam, J, “Roskam, J, “Airplane Design, Part VI: Preliminary Calculation ofAerodynamic, Thrust, and Power Characteristics,” DARcorporation,Lawrence, Kansas, 2004

[4] Howe, D., “Aircraft Loading and Structural Layout,” AIAA Educational Series,Virgina, 2004


g ,

[5] Lomax, T., “Structural Loads Analysis for Commercial Transport Aircraft: Theroyand Practice,” AIAA Educational Series, Virgina, 1995

Critical Load Cases for Transport Aircraft:


For each load condition the structures team requires the total aerodynamic forces and pressure distribution for……

Classical Critical Load Cases, Nui(1)

Load case Description

Pilot Induced Maneuvering and System malfunctions

Combination stabilizer-elevator maneuvers

Longitudinal maneuvers which can load the LcCE and main wing.

Aileron and/or spoiler Lateral maneuvers which apply an asymmetric loading condition on the pmaneuvers

pp y y gwing.

Rudder maneuvers Directional maneuvers which the rudder applies loads to the vertical stabilizer and fuselage

Atmospheric Turbulence Power spectral approach and/or Discrete gust approach

Landing Landing gear and aerodynamic loads encountered during landing

Ground handling

FAR ground handling Loads encountered during taxing maneuvers

Rotational Taxing Loads encountered during taxing maneuvers

Rotational ground maneuver Loads encountered take-off rotation

Jacking Loads jacking of aircraft for maintenance purposes

Fail-safe and breakaway design


4.6 Assignment


Assignment Aerodynamic for Performance:

• Produce a low speed 2-D drag polar and lift curve for the GIIIBL45 airfoil

• Produce the 3-D drag polar for T-O, Climb, Cruise, and Approach.

• Produce a Digital DATCOM model of the Citation and compile the long range cruise drag polar and Lift curve

The report should include,

1. Quick summary of the capability and limitations of X-FOIL and Digital DATCOM

2. 2-D drag polar from X-FOIL3. The 3-D drag polar and lift curves from Digital DATCOM,

AVL and hand-Calculations• Plot 2-D and 3-D CL vs. CD, CL vs. AOA and L/D vs. CL

• Tabulate L/Dmax and CL and CD and L/Dmax


4.6 Assignment


Assignment Aerodynamic for Stability and Control:

• Construct a Digital DATCOM and AVL model of the Citation X

• Trim the aircraft for Cruise and Approach (with appropriate configuration settings).

• Report the resulting stability and control derivatives


1 Quick summary of the capability and limitations of AVL1. Quick summary of the capability and limitations of AVL and Digital DATCOM

2. Brief description of the trim method used3. Table summarizing the stability and control derivatives

during Cruise and Approach 4. Comparison of Digital DATCOM and AVL results


4.6 Assignment


Assignment Structural Loads:

• Build an AVL Model Citation X• Trim the aircraft during Cruise, and Approach by

constraining the AOA to match the CL required. • Produce the lift and pressure distribution for the lifting

surfaces.


1. Quick summary of the capabilities and limitations of the vortex lattice code AVLvortex lattice code AVL

2. 3-D wire-frame drawing of the Citation X model3. AVL Lift distribution and pressure distribution plot for

Cruise and Approach

Due: Update Report Friday 11-20-09 at 5:00 pm Final Report Friday 12-4-09 at 5:00 pm


mae 4350 - lab 4 aerodynamics

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