AERODYNAMIC ANALYSIS

OF THE

CARBON DRAGON AIRFOILS

PART ONE

 

 

 

About the author:

Alejandro Ramirez-Pineiro (30) has the degree of Bachelor of Science in Mechanical Engineering, earned at the Universidad Tecnologica Metropolitana of Santiago, Chile. At this moments he is working as a Composite Materials Engineer at ENAER, building a 100% composite JAR 23 airplane (visit www.euro-enaer.com ). Also he is building a modified Carbon Dragon, planned to fly during the first quarter of 2000. You can reach him at alejo@netline.cl .

 

 

Introduction:

The Carbon Dragon is an ultralight, high performance glider.  This glider was conceived by Jim Maupin who also designed the famous Woodstock sailplane and the Windrose motorglider.  With encouragement from Irv Culver, who promised to "run the numbers", Jim Maupin finished the design and built the prototype.  As a result, Jim Maupin with an assist from co-designer Irv Culver, produced an exciting sailplane.

 

 

Scope of the study:

The proposal of this study is getting aerodynamic data that can be used for the structural modifications of the original Carbon Dragon. This  data will be used also on  the study of the longitudinal static stability of the modified glider. All the aerodynamic values presented here can be useful to actual or future builders for helping their own studies.

            For preliminary design, the values presented in this report are excelent, but must to keep in mind that the data presented here was developed by a computer software, so, to get more “real” values, the reader will have to do wind tunnel testing.

 

 

Getting the coordinates:

            To get the coordinates of the wing root and wing tip airfoils, the original drawings of the ribs number 1 and number 13 where divided on 88 stations on the root and 58 stations on the tip. The mesures where taken with an accuracy of  + - 0.5 mm. The job was done by my building partner Miguel Eyquem.

            With the coordinates obtained in milimiters, the next step was to convert them in fractions of chord. For this task I used an electronic spreadsheet.

 

 

 

Airfoil Coordinates of the Wing Root Airfoil

(values on fraction of chord)

 

Coord

x

Yu

Yl

 

Coord

X

Yu

Yl

0

0

0

0

 

51

0.361367

0.116951

-0.07819

1

0.001314

0.00657

-0.00591

 

52

0.374507

0.115966

-0.07819

2

0.003285

0.010184

-0.00828

 

53

0.387648

0.114652

-0.07806

3

0.00657

0.014783

-0.0115

 

54

0.400788

0.113338

-0.07779

4

0.009855

0.018725

-0.0138

 

55

0.413929

0.111695

-0.0772

5

0.013141

0.022339

-0.01577

 

56

0.42707

0.110381

-0.07668

6

0.016426

0.026281

-0.01774

 

57

0.44021

0.108541

-0.07615

7

0.019711

0.029566

-0.01971

 

58

0.453351

0.106767

-0.0753

8

0.022996

0.032852

-0.02102

 

59

0.466491

0.105125

-0.07424

9

0.026281

0.035808

-0.02234

 

60

0.479632

0.103154

-0.07326

10

0.029566

0.038765

-0.02431

 

61

0.492773

0.101183

-0.07194

11

0.032852

0.041393

-0.02562

 

62

0.505913

0.099212

-0.0707

12

0.036137

0.044678

-0.02727

 

63

0.519054

0.096912

-0.06932

13

0.039422

0.047306

-0.02858

 

64

0.532194

0.094941

-0.06754

14

0.042707

0.049934

-0.02989

 

65

0.545335

0.092576

-0.06597

15

0.045992

0.052562

-0.03121

 

66

0.565046

0.08883

-0.06334

16

0.049277

0.054534

-0.03252

 

67

0.584757

0.085414

-0.06064

17

0.052562

0.056505

-0.03351

 

68

0.604468

0.08134

-0.05782

18

0.055848

0.059133

-0.03463

 

69

0.624179

0.076938

-0.05519

19

0.059133

0.061104

-0.03581

 

70

0.64389

0.07293

-0.05256

20

0.062418

0.063403

-0.03679

 

71

0.663601

0.06866

-0.04961

21

0.065703

0.065703

-0.03784

 

72

0.683311

0.064389

-0.04652

22

0.072273

0.068988

-0.04008

 

73

0.703022

0.05979

-0.04336

23

0.078844

0.072273

-0.04172

 

74

0.722733

0.055191

-0.04074

24

0.085414

0.075558

-0.04382

 

75

0.742444

0.050591

-0.03752

25

0.091984

0.078844

-0.04534

 

76

0.762155

0.047306

-0.03476

26

0.098555

0.0818

-0.04731

 

77

0.781866

0.043364

-0.03187

27

0.105125

0.084757

-0.04901

 

78

0.801577

0.039093

-0.02891

28

0.111695

0.087385

-0.05099

 

79

0.821288

0.035085

-0.02602

29

0.118265

0.090013

-0.05256

 

80

0.840999

0.031012

-0.023

30

0.124836

0.092313

-0.05414

 

81

0.86071

0.026938

-0.0203

31

0.131406

0.094415

-0.05565

 

82

0.88042

0.022996

-0.01537

32

0.137976

0.096583

-0.05696

 

83

0.900131

0.018922

-0.01445

33

0.144547

0.098555

-0.05848

 

84

0.919842

0.01498

-0.0113

34

0.151117

0.100657

-0.05999

 

85

0.939553

0.010841

-0.00821

35

0.157687

0.102497

-0.0611

 

86

0.959264

0.006899

-0.00512

36

0.164258

0.104205

-0.06242

 

87

0.978975

0.002957

-0.0021

37

0.177398

0.107753

-0.06472

 

88

1

0

0

38

0.190539

0.11071

-0.06669

 

 

 

 

 

39

0.203679

0.113009

-0.06866

 

 

 

 

 

40

0.21682

0.115375

-0.07063

 

 

 

 

 

41

0.229961

0.116754

-0.07194

 

 

 

 

 

42

0.243101

0.118265

-0.07293

 

 

 

 

 

43

0.256242

0.119054

-0.07392

 

 

 

 

 

44

0.269382

0.119842

-0.0749

 

 

 

 

 

45

0.282523

0.119908

-0.07589

 

 

 

 

 

46

0.295664

0.11958

-0.07654

 

 

 

 

 

47

0.308804

0.119251

-0.07727

 

 

 

 

 

48

0.321945

0.119054

-0.07753

 

 

 

 

 

49

0.335085

0.118594

-0.07786

 

 

 

 

 

50

0.348226

0.117937

-0.07819

 

 

 

 

 

 

 

 

 

 

 


Image 1. Carbon Dragon Wing Root Airfoil

 

 

 

 


Airfoil Coordinates of the Wing Tip Airfoil

(values on fraction of chord)

 

 

Coord

X

Yu

Yl

 

Coord

X

Yu

Yl

0

0

0

0

 

31

0.363278

0.100036

-0.0349

1

0.00179

0.011632

-0.01163

 

32

0.381174

0.098425

-0.03364

2

0.005369

0.020401

-0.01825

 

33

0.399069

0.096994

-0.03239

3

0.010737

0.027738

-0.02523

 

34

0.416965

0.095204

-0.03185

4

0.017895

0.034896

-0.03114

 

35

0.43486

0.093057

-0.03042

5

0.025054

0.040444

-0.0349

 

36

0.452756

0.091267

-0.02953

6

0.032212

0.046528

-0.03758

 

37

0.470651

0.089298

-0.02845

7

0.03937

0.051897

-0.03937

 

38

0.488547

0.086793

-0.02756

8

0.048318

0.057266

-0.04116

 

39

0.506442

0.084109

-0.02649

9

0.057266

0.063529

-0.04295

 

40

0.524338

0.081424

-0.02523

10

0.066213

0.068003

-0.04384

 

41

0.542233

0.07874

-0.02416

11

0.075161

0.072656

-0.04474

 

42

0.560129

0.077666

-0.02309

12

0.085898

0.076951

-0.04545

 

43

0.578024

0.072656

-0.02219

13

0.094846

0.08053

-0.04563

 

44

0.59592

0.069792

-0.02112

14

0.103794

0.083751

-0.04581

 

45

0.613815

0.066571

-0.02004

15

0.112742

0.086256

-0.04563

 

46

0.631711

0.063529

-0.01915

16

0.121689

0.089298

-0.04617

 

47

0.649606

0.060845

-0.0179

17

0.130637

0.091267

-0.04635

 

48

0.667502

0.057266

-0.01646

18

0.139585

0.093057

-0.04653

 

49

0.685397

0.054581

-0.01611

19

0.148533

0.094846

-0.04635

 

50

0.721188

0.048318

-0.01432

20

0.166428

0.098067

-0.04384

 

51

0.756979

0.041696

-0.01217

21

0.184324

0.100215

-0.04295

 

52

0.79277

0.035791

-0.01056

22

0.202219

0.101467

-0.04259

 

53

0.828561

0.029707

-0.00823

23

0.220115

0.102362

-0.0417

 

54

0.864352

0.023264

-0.00626

24

0.23801

0.103078

-0.0408

 

55

0.900143

0.01718

-0.00447

25

0.255906

0.103615

-0.03991

 

56

0.935934

0.010916

-0.00268

26

0.273801

0.103794

-0.03919

 

57

0.971725

0.004295

-0.00089

27

0.291696

0.103794

-0.03812

 

58

1

0

0

28

0.309592

0.102899

-0.0374

 

 

 

 

 

29

0.327487

0.102004

-0.03669

 

 

 

 

 

30

0.345383

0.10111

-0.03579

 

 

 

 

 

 

 

 

 

 

 

 


Image 2. Carbon Dragon Wing Tip Airfoil.

 

 

 

 

 


Using the software:

            The software used was Airfoil Analysis Geometric Module (to get more information and/or demos visit http://airfanalysis.hypermart.net ) developed by two Engineers at Italy.

            The data was imported from a text file into the Coorditate Editor to make easier the data imput to the software.

            Using both airfoils (root and tip) was possible to get by interpolation the mean geometric chord airfoil, slightly different from the Mean Aerodynamic Chord, for a simply, linearly tapered wing. For this task I used the Airfoil Mix Window and set the mix proportion to 50 %. See the Mean Geometric Chord coordinates table.

            With the coordinates of the root, tip and Mean Geometric Chord airfoils the software was able to give the following information:

 

            1.- Geometric data of the airfoils

            2.- Estimated aerodynamic data, including:

·        Angle of atack for zero lift, a 0

·        Pitching moment at zero lift, Cm0

·        Lift slope, a0

 

 

Airfoil Coordinates of the Wing Mean Geometric Chord Airfoil

(values on fraction of chord)

 

Coord

x

Yu

Yl

0

0

0

0

1

0.0016

0.0092

-0.0087

2

0.0065

0.0184

-0.0156

3

0.0145

0.028

-0.0227

4

0.0257

0.0381

-0.0286

5

0.04

0.05

-0.0342

6

0.0573

0.0618

-0.039

7

0.0774

0.0726

-0.0431

8

0.1003

0.0826

-0.0468

9

0.1257

0.0915

-0.0503

10

0.1536

0.0986

-0.0531

11

0.1838

0.1047

-0.0543

12

0.216

0.1087

-0.0562

13

0.25

0.1111

-0.0568

14

0.2857

0.1119

-0.0573

15

0.3227

0.1107

-0.0572

16

0.3609

0.1086

-0.0566

17

0.4

0.1052

-0.0551

18

0.4397

0.1006

-0.0531

19

0.4799

0.0956

-0.0507

20

0.5201

0.0894

-0.0473

21

0.5603

0.0837

-0.0435

22

0.6

0.0757

-0.0396

23

0.6391

0.0681

-0.036

24

0.6773

0.0607

-0.0318

25

0.7143

0.0534

-0.0283

26

0.75

0.0461

-0.0245

27

0.784

0.0401

-0.0213

28

0.8162

0.034

-0.0179

29

0.8464

0.0282

-0.0148

30

0.8743

0.0229

-0.0113

31

0.8997

0.0181

-0.0095

32

0.9226

0.0138

-0.0071

33

0.9427

0.0099

-0.005

34

0.96

0.0066

-0.0032

35

0.9743

0.0039

-0.0018

36

0.9855

0.0019

-0.0008

37

0.9935

0.0007

-0.0003

38

0.9984

0.0002

0

39

1

0

0

 

 

 

 

Image 3. Carbon Dragon Mean Geometric Chord.

 

 

 

 

 

Geometric Data of the Carbon Dragon Airfoils

 

 

 

Wing root airfoil

Mean Geometric  Chord airfoil

Wing tip airfoil

Maximum thickness

19.67 %

16.91 %

14.42 %

Position max. thickness

31.30 %

28.43 %

20.87 %

Maximum camber

2.27 %

2.74 %

3.29 %

Position max. camber

24.55 %

27.30 %

29.57 %

Leading edge radius

0.68 %

1.67 %

2.19 %

Trailing edge angle

17.23 degrees

14.44 degrees

11.63 degrees

Geometric centroid in X

41.43 %

40.12 %

38.34 %

Geometric centroid in Y

1.56 %

2.05 %

2.49 %

 

 

 

 

Estimated Aerodynamic Data of the Carbon Dragon Airfoils

 

 

Wing root airfoil

Mean Geometric  Chord airfoil

Wing tip airfoil

a 0 [degrees]

-1.27

-1.92

-2.57

Cm0

-0.0158

-0.0357

-0.0557

Lift curve slope efficiency factor m

1.0153

1.0132

1.0112

Lift slope, a0 [1/degree]

0.111

0.111

0.110

 

 

Observations regarding the geometric characteristics of the airfoils:

 

 

a) The increasing camber toward the tip should be due to a research of higher maximum lift coefficients to avoid tip stalling.

 

b) Progressively moving forward the maximum center position should be due to obtain earlier transition on the tip section and minimize Reynolds number effects.

 

c) The increase of the L.E. radius help to soften tip stall.

 

 

For the eager builder/designer:

 

Some remarks about the “estimated” aerodynamic data presented in AIRFOIL ANALYSIS – GEOMETRIC MODULE: The estimates of zero lift angle and zero lift quarter chord moment coefficient are obtained by the Pankurst’s method (“Theory of Wing Sections”, by Abbott and von Doenhoff) … relying just upon the knowledge of the mean line shape: it’s a sort of Gaussian quadrature and it should had been developed analytically, so no numeric is still present at this stage. The method is surprisingly good, relating to its simplicity, as you own will experience when comparing estimated to calculated data (by the numerical panel method or comparison with experiments). In fact, the flow quality on the airfoil surfaces is still “good” in the proximity of the zero lift angle (that is usually small) for conventional, unflapped airfoils … this explains why these estimates are incredibly useful even if absolutely “a priori” of any aerodynamic analysis as properly defined.

The lift curve slope efficiency factor m (the ratio of the potential flow slope to the theoretical value of 2p) is estimated according to R. Eppler (“Airfoil Design and Data”, Springer Verlag) but it’s usefulness is usually driven off by viscous effects … so practically it doesn’t matter anymore when viscous analysis are on hand.

 

References:

 

Airfoil Analysis – Geometric Module

http://airfanalysis.hypermart.net

 

Engineer Luca Cistriani

http://airfanalysis.hypermart.net

 

Carbon Dragon Technical Website - by S. Steve Adkins

http://www.isd.net/sadkins/builders.htm

 

Jim Maupin Ltd.

www.jcpress.com/JMaupinLtd/carbon.htm

 

 

 

            The part II of this study will include: