Effects of Mach cruise number on conventional civil jet aircraft sizing

BACHELOR THESIS IN AERONAUTICAL ENGINEERING15 CREDITS, BASIC LEVEL 300

School of Innovation, Design and Engineering

Effects of Mach cruise number on conventional

civil jet aircraft sizing

Author: Niklas Bergman Report code: MDH.IDT.FLYG.0216.2009.GN300.15HP.Ae

ii

ABSTRACT

This thesis work was executed at Swift Engineering Incorporated located in San

Clemente, California during spring in 2009. Placement supervisor from Swift was Mark Page

and advisor and examiner from the Division of future products at Mälardalen University,

Sweden was Gustaf Enebog.

The objective with this thesis work was to examine the effects of fitness ratio, lift over

drag, lift coefficient at cruise, winglet span, wing sweep angle, wing aspect ratio, wing area

and weights with respect to Mach number for a conventional business jet capable of 18

passengers. The cruise speed study range from Mach 0.88 to 0.99.

The Excel based conceptual design tool Jetsizer 2008c was used to make four models

with similar configuration and mission but with different cruise Mach numbers.

A new Jetsizer module was then created to handle a modification process where the

models are optimized for their speed and configuration. The result in this report gives

guidelines for the needed values when creating an initial CFD model for this type of airplane.

Date: 31 July 2009

Carried out at: Swift Engineering Inc.

Advisor at MDH: Gustaf Enebog

Advisor at Swift Engineering Inc: Mark Page

Examinator: Gustaf Enebog

iii

NOMENCLATURE

Abbreviation Declaration

AR Wing Aspect Ratio

c Chord of wing

CD Drag coefficient

CD,i Induced drag coefficient

CL CRZ Lift coefficient at cruise speed

CFD Comptional fluid dynamics

ICA Initial cruise altitude

ISA International Standard Atmosphere

L/D Lift over Drag

M Mach number

MCRZ Mach number at cruise

MLW Maximum Landing Weight

MTOW Maximum TakeOff Weight

ROC Rate Of Climb

SL Sea Level

SLST Sea Level Standard Thrust

iv

TABLE OF CONTENTS

Chapter 1 INTRODUCTION 1

1.1 Background ..........................................................................................................1

1.2 Objective ..............................................................................................................1

1.3 Problem formulation .......................................................................................... 2

1.4 Limitations .......................................................................................................... 2

Chapter 2 METHOD 3

2.1 Mission ................................................................................................................. 3

2.2 Sizing tool............................................................................................................ 4

Sizing tool modification ............................................................................................................ 4

2.3 Baseline model .................................................................................................... 5

2.4 Mach study models ............................................................................................. 6

2.5 Geometry modification tool ................................................................................ 7

Chapter 3 RESULTS 9

Winglet height study ............................................................................................................... 12

Chapter 4 DISCUSSION 14

Chapter 5 FUTURE WORK 15

Chapter 6 ACKNOWLEDGEMENTS 16

Chapter 7 REFERENCES 17

Appendix A Baseline model for M=0.88

Appendix B Optimized model for M=0.88

Appendix C Optimized model for M=0.92

Appendix D Optimized model for M=0.96

Appendix E Optimized model for M=0.99

Appendix F Program code for the modification tool

1

Chapter 1

INTRODUCTION

1.1 Background

Swift was consulted by a customer to design and build a civil jet aircraft based on a

special developed wing structure. The requirements where similar to the daily operating civil

jets but with a preferably higher cruise speed near Mach one. In order to design this airplane

an initial study was needed to investigate the relations of aircraft geometry and the

aerodynamic effects on an airplane with cruise speed near Mach one.

This study was needed to be able to generate a conventional baseline which allows Swift

to develop the initial CFD geometry for this project.

1.2 Objective

The purpose of this study is to demonstrate the effects of Mach number ranging from

0.88 to 0.99 for a conventional business jet in respect to

• fitness ratio

• lift over drag

• lift coefficient at cruise

• winglet span

• wing sweep angle

• wing aspect ratio

By creating a number of aircraft models with the same configuration but optimized for

different speeds a study can be made to compare the geometry variation due to the Mach

number. This result will give guidelines for the geometry when making CFD models for a

certain cruise speed.

2

1.3 Problem formulation

First a mission has to be defined in order to be able to define a design and to have a

meaningful study.

A tool has to be chosen to use in the aircraft sizing stage in order to generate different

models that can be compared in a Mach cruise study. It is recommended to use Jetsizer 2008c

for this study but it is maybe necessary to implement some new modules to estimate the

transonic wave drag since Jetsizer does not take that in count.

Another tool has to be created to handle the input values in the optimization process and

to display the results from each iterated step.

1.4 Limitations

This survey is only aimed at generating guidelines for a future conceptual study. For this

reason there are some limitations.

• Engine inlet effects have not been taken account of.

• Limited to the size and configuration of a conventional civil jet aircraft capable of

caring 18 passengers.

3

Chapter 2

METHOD

When design an aircraft from scratch you usually start with determine the mission

profile. In the early phase of this project there was no specific mission. The only given

specification was that it should be a civil jet aircraft with competitive performance of modern

aircrafts in the same category.

The first step was to do a basic investigation of already constructed and functional civil

jet aircrafts and take their mission and performance as guidelines for this project.

In order to calibrate and to see how well Jetsizer correlate its calculations a model of the

Gulfstream G650 was entered in Jetsizer with a set mission.

2.1 Mission

The mission for the models was determined by looking at other aircrafts. A summary of

airplane performance was found in Aviation week & space technology1. It was decided to use

the Gulfstream G650 as a calibration baseline. The G650 has a very high cruise speed

together with a very long range as shown in table 1 below. The results made it to be an

interesting baseline for the initial study in Jetsizer.

Table 1: Mission for Gulfstream G650

1 Aviation week & space technology, Aerospace source book, January 26, 2009

Maximum Range 7000 nm (Mach 0.85)

Normal Cruise Mach 0.85 / 488 ktas

High Speed Cruise Mach 0.90 / 516 ktas

Takeoff Distance (SL, ISA, MTOW) 6000 ft

Landing Distance (SL, ISA, MLW) 3000 ft

Initial Cruise Altitude 41000 ft

Maximum Takeoff Weight 99600 lb

Basic Operating Weight 54000 lb

Maximum Fuel Weight 44200 lb

Passengers 11 - 18

Rated Takeoff Thrust 16100 lb

4

The final mission and requirements for the Mach study was determined from the G650.

An increase of the Mach number will affect some of the mission variables there for are only a

few clarified. Table 2 shows the mission for the models in the mach study.

Table 2: Mission for the increased mach study model

2.2 Sizing tool

Jetsizer2008c was used as the sizing tool for this research. It is user friendly and it is

possible to remodel the program for suiting the study since the source code is open. It is

capable of the most important calculations in the conceptual design stage of general

configured passenger jets. Since Jetsizer is based on Excel, it is very easy for the user to work

through another Excel documents and by a macro import and export data for analysis.

Sizing tool modification

The wing area in Jetsizer is calculated in a conventional manner including the whole

wing span through the fuselage. This gives an incorrect geometry when entering a specified

wing area that does not involve the middle section through the fuselage. When sizing an

airplane it makes a good assumption to let the configured wing for a conventional design pass

all the way through the fuselage to obtain elliptical lift distribution. When calculating the

drag, this section through the fuselage gives a drag rise that Jetsizer does not neglect. This

drag rise was left in the calculations to be canceled by other drag distributing sources that

Jetsizer does not recognize e.g. interference between wing and fuselage.

Business jet passengers demand a higher comfort than a regular commercial airliner

offers. There for a comfort factor was added to simulate the assumed weight increased for

cabin equipment. The comfort factor works as a multiplier to the passenger weight.

Jetsizer 2008c lacks the ability to handle cruise speed near Mach one. A module was

implemented to handle compressibility interference between components. This module is

based on two papers from NACA, “Examples of the Applications of the Transonic and

Supersonic Area Rules to the Prediction of Wave Drag” (references number 7) and “A Study

of the Zero-Lift Drag-Rise Characteristics of Wing-Body Combinations near the Speed of

Sound” (references number 8). The results and the equations are implemented in Jetsizer to

simulate the transonic wave drag due to volume.

Maximum Range @ ICA & Mcrz 7000 nm

Takeoff Distance (SL, ISA, MTOW) 11000 ft

Landing Distance (SL, ISA, MLW) 5000 ft

Initial Cruise Altitude 41000 ft

Passengers 18

Mission revenue cargo 2630

% Mission fuel @ MLW 10%

5

2.3 Baseline model

The first model that was made in Jetsizer was based on data from the Gulfstream G650.

The unpublished geometry was measured from a 3-view picture and entered in Jetsizer. The

model was not supposed to be identical with the G650 but similar enough to see how well

Jetsizer correlates with the published performance.

Table 3: Geometry of Gulfstream G650 that was used in the baseline model.

Wing

AR 7.73

Area 1283 ft2

Axial position 366 ft

Dihedral 3 deg

Span 99.58 ft

Sweep @ c/4 33 deg

Horizontal stabilizer

AR 4.84

Dihedral 0 deg

Sweep @ c/4 33 deg

Taper ratio 0.47

Vertical stabilizer

AR 1.07

Sweep @ c/4 30

Taper ratio 0.95

Fuselage

Total length 1043.3 in

Cabin length 643 in

Cabin width 102 in

Cabin height 77 in

The generated results were at first

not satisfying but after some

modifications in Jetsizer and to the

model it correlated with the G650. The

modifications were; inserting the

comfort factor in Jetsizer, changing

structural materials to correlate the

empty weight and modifying wing area

to meet the performance.

The baseline model was later used

in the mach study as the model with the

lowest cruise speed, Mach 0.88.

The complete baseline model is found in

appendix A.

Figure 1: Baseline model from Jetsizer

6

2.4 Mach study models

Four different models were entered in Jetsizer with a conventional civil jet configuration

and with individually determined mach cruise numbers ranging from 0.88 to 0.99. The

purpose was to have four optimized models for each cruise speed with similar configuration.

In that way it is possible to see the effects of geometric related to the cruise speed.

The mission used for the models is determined in chapter 2.2.

To keep the models comparable regarding to cabin size the middle fuselage section

diameter was kept constant as well as the cabin floor area. The upper and lower lobe diameter

was set to 102 inches and with a separation vertically of 3 inches. The cabin floor length was

set to 767 inches. These measurements are based on the G650 and defined by the first

baseline model.

Figure 2: Fuselage cross

section of the models.

The four models were all generated from the baseline model and then converged with a

defined cruise Mach number. Each model works as a baseline for the individual Mach

number. The models were optimized by using the geometry modification tool. By taking the

best values from the iterating process a new model for each study was generated. Data from

the final optimized models can be found in Appendix B to E.

7

2.5 Geometry modification tool

To investigate the relations between the geometry and Mach number a tool was needed

to handle the input data and the communication with the sizing tool and to display the

calculated results. This tool was made in excel with a macro to control the calculations and

the communication with Jetsizer. The specific models are designed in Jetsizer after a

defined cruise Mach number which allows the modification tool to only look at one model

and Mach number at a time. By combining the results from several studies it is possible to

make a trend of the Mach number effects.

The application works directly under Jetsizer were the user is allowed to select a range

of values for a specified study. For each step Jetsizer is told to modify and converge to the

model for the best result. The results are given back to the application and are shown in a

table.

Figure 3: The loop of the geometry modification tool.

8

The application is limited to handle only six different studies for each model (AR, wing

sweep, winglet span, fuselage length, bypass ratio and SLST) which was decided to be

enough for this study. Figure 4 shows the interface were the user is allowed to enter a

number or a series of numbers in the cells to the right of AR, Sweep, Winglet span,

Fuselage length, Bypass ratio or SLST.

In the top left corner the user is allowed to enter the Jetsizer file name that wants to be

used. The only criterion is that both the active Jetsizer file and this tool have to be saved in

the same map. The archive is directly connected to which archived model in Jetsizer the

user wants to use.

By pressing one of the buttons to the left for each study respectively the program will

start the calculation on the active cell and continue as long as there is a number written on

that same row. The results will be shown in respectively column below where it is easy to

see the trend or copy the values to make a plot.

Figure 4: The interface of the geometry modification tool.

9

Chapter 3

RESULTS

The project resulted firstly in a tool that can be used with Jetsizer to optimize different

geometries. Although it is limited to only a few different modification options it can easily be

extended for other studies of interest by copying the code and entering new addresses in the

links to Jetsizer.

3.1 Mach cruise study

The results from the optimization tool is gathered and summarized in some plots. These

shows the trend lines for the increasing Mach number and the modifications needed to the

models.

0

20 000

40 000

60 000

80 000

100 000

120 000

0,86 0,88 0,9 0,92 0,94 0,96 0,98 1

lbs

Mach number

Weights

TOGW

OEW

FB

Figure 5. Trend line of weight versus Mach number.

10

The total weight is increasing a lot due to higher speed as shown in figure 5. This is an

effect of the acting transonic wave drag when traveling in high speeds. The rise of the

operating empty weight is an effect of the increasing wing area that is needed for high speeds.

It can be seen in figure 6 that the aspect ratio is descending when the Mach number goes

up. By lowering the AR the fuel volume in the wings increases and the weight of the wing

structure decreases. ROC decrease and TOFL gets longer. This is an effect of CD,i that is

inversely proportional to the aspect ratio (equation 3.1) which then affects the total drag

(equation 3.2). With a higher Aspect Ratio the induced coefficient of drag will decrease.

��,� =��

��

�� = � + ��,�

The increasing wing area with rising Mach number is an effect of lower aspect ratio and

the larger fuel capacity needed for higher speed.

800

850

900

950

1 000

1 050

1 100

1 150

1 200

1 250

0

1

2

3

4

5

6

7

8

9

0,86 0,88 0,9 0,92 0,94 0,96 0,98 1

Win

g A

rea

, ft

2

AR

Mach number

Wing Aspect Ratio & Area

AR

Sw

Figure 6. Trend line of AR and Sw versus Mach number.

(3.1)

(3.2)

11

In figure 7 it is shown the result of wing sweep angle when converging to the lowest CD

possible. With a grater sweep the total coefficient of drag is reduced and there for ROC is

increased. Greater sweep gives a heavier wing structure and a lower fuel capacity which is

there for compensated with a larger wing area.

41

42

43

44

45

46

47

48

49

50

0,86 0,88 0,9 0,92 0,94 0,96 0,98 1

Sw

ee

p a

ng

le

Mach number

Wing Sweep

Swe…

Figure 7. Trend line of Wing sweep versus Mach number.

12

The reason for the dropping Lift over Drag curve in figure 8 is generally because the drag

gets significantly higher when it closes to Mach one. Also the increase of weight due to

higher speed affects the lift directly. The CLcrz curve is mainly effected by the Weight devided

by the wing area which both increases with the higher speeds.

Winglet height study The winglet span (or height) is proportional to the vortex drag and is also affecting the

weight of the wing structure and fuel quantity.

When extending the winglets the vortex drag decreases and the weight of the wing

increases this result in a loss of total fuel because of the change of the wing structure and a

lower CD. When adding a winglet in Jetsizer it affects the CD and when extending the winglet

you will get an even lower CD.

In the winglet height study it was discovered that Jetsizer uses a simplified formula for

the winglet calculations. This results in an endless decrease of CD when the winglet height

goes towards infinity. This result rejected the winglet span study.

0,000

0,050

0,100

0,150

0,200

0,250

0,300

0,350

0,400

0,450

0,500

13,50

14,00

14,50

15,00

15,50

16,00

16,50

17,00

0,86 0,88 0,9 0,92 0,94 0,96 0,98 1

CLc

rz

L/D

Mach number

L/D & CLcrz

L/D

CLcrz

Figure 8. Trend line of L/D and CLcrz versus Mach number.

13

Figure 9 shows the Fuselage length trend line, same as fuselage length to thickness ratio in

this study because the thickness does not vary between the models. This curve is only affected

by the transonic wave drag in this study.

1 000

1 050

1 100

1 150

1 200

1 250

1 300

1 350

0,86 0,88 0,9 0,92 0,94 0,96 0,98 1

Len

gth

, in

ch

Mach number

Fuselage length

Fuselage length

Figure 9. Trend line of fuselage length versus Mach number

14

Chapter 4

DISCUSSION

First of all, this study is only meant to give guidelines for the initial input values of a

future conceptual study on this type of aircraft. The characteristics of the results are known

from fundamental aeronautical equations but the numerical values related to the speed and

geometry is individual for each aircraft. The result of the plots shows expected shapes in the

higher Mach region and by comparing other airplanes in the same category it shows similar

results.

For M=0.88 and M= 0.92 the results show no or only a slight change in the performance

which should need a more accurate investigation. This could be an effect of converge the

models to the lowest CD. Probably should the M=0.88 study be rejected because that case is

not optimized in perspective of the compressibility interference module. Instead it is made

directly from the baseline of G650.

15

Chapter 5

FUTURE WORK

• Improve the modification tool with more input variables and include it in Jetsizer.

• Improve the winglet formula in Jetsizer.

• Improve the compressibility interference module in Jetsizer.

16

Chapter 6

ACKNOWLEDGEMENTS

I want to thank my supervisor Mark Page at Swift Engineering for arranging this thesis

work and made it possible for me to join them at Swift for 10 weeks. I have to thank Mark for

all his help and for his imaginativeness that usually solves even the toughest problems.

Thanks to all the guys at Swift; Ed Smetak, for helping me arrange a living during the

stay and driving me to work every day and letting me experience the life in California. Any

Luo, for supervising me in this project and helping me with everything associated with

computers and aerodynamics. John Winkler, for your experienced help with aeronautical

topics and for showing me around in the mountainous Californian run track.

Thanks to Chris Thompson for letting a complete stranger from Sweden to occupy one of

your wonderful bedrooms in the magnificent beach house.

Special thank to my supervisor Gustaf Enebog at MDH, Sweden, for helping me with the

formal elements of this thesis work and for the contacts who gave me this opportunity.

I finally want to thank the international unit at MDH for financing my time abroad.

Niklas Bergman

17

Chapter 7

REFERENCES

1. Dr. Jan Roskam. Airplane Design Part I,”Preliminary Sizing of Airplanes”, DARcorporation

2005, ISBN-13: 978-1-884885-42-6

2. Dr. Jan Roskam. Airplane Design Part II, ”Preliminary Configuration Design and Integration

of the Propulsion System” ,DARcorporation 2005, ISBN: 1-884885-43-8

3. Dr. Jan Roskam. Airplane Design Part III, ”Layout Design of Cockpit, Fuselage, Wing and

Empennage: Cutaways and Inboard Profiles”, DARcorporation 2005, ISBN: 1-884885-56-x

4. Dr. Jan Roskam. Airplane Design Part VI, ”Preliminary Calculations of Aerodynamic, Thrust

and Power Characteristics”, DARcorporation 2005, ISBN-13: 978-1-884885-

52-5

5. Dr. Jan Roskam. Airplane Design Part VII, ”Determination of Stability, Control and

Performance Characteristics: FAR and Military Requirements”, DARcorporation 2005, ISBN-

13: 978-1-884885-54-9

6. Aviation week & space technology, Aerospace source book, January 26, 2009

7. Robert L. Nelson Clement J. Welsh. “Examples of the Applications of the Transonic and

Supersonic Area Rules to the Prediction of Wave Drag”, NACA RM L56D11, Langley

Aeronautical Laboratory, Langley Field, Va., 20 March 1957.

8. Whitcomb, Richard T. “A Study of the Zero-Lift Drag-Rise Characteristics of Wing-Body

Combinations Near the Speed of Sound”, NACA RM L52H08, Langley Aeronautical

Laboratory, Langley Field, Va., 3 September 1952.

9. John D. Anderson, Jr. Fundamentals of Aerodynamics, Mc Graw Hill 2007, ISBN: 007-

125408-0

18

A. APPENDIX

Baseline model cruise speed Mach 0.88.

Fuselage & Materials

Ncrew = 2 n.d. Typ 3 class

Nattendants = 2 n.d. Natt =1

Diameter (avg), Dfus = 8,6 ft Dfus =6,1

Total Length, Lfus = 87 ft Lfus =59

# Coach Seats Abreast = 2 n.d. -

Number of Aisles = 1 n.d. -

Fineness Ratio (L/Dia)fus = 10,0 n.d.

LH2:TankWt/FuelWt | InsulatnWt/Area = 0,0 0,0

n.d. | lbs/ft2

Wing Materials = COMP "AL", "ALLI", "COMP"

Fuselage Materials = COMP "AL", "ALLI", "COMP"

Empennage Materials = COMP "AL", "ALLI", "COMP"

Fuel Type = Kerosene "Kerosene", "LH2"

Propulsion

SLST per engine = 14 600 lbs 2 Engines

Neng = 2 n.d.

Eng Location = F

"W"ing, "F"uselage

Engine Cert Date & EIS = 1980 year

Bypass Ratio (BPR) = 4,4 n.d. **TFAN**

SexposedPylon = 86,0 sq-ft per pylon

Fuel Price = 1,1 $/USgal

Geometry (Trap) Wing Horiz Vert Planform Area, S = 950,0 207,2 103,8

Span, b = 87,2 26,9 10,9

19

Mean Aero Chord, MAC = 12,2 8,0 9,9

Croot = 17,4 10,5 12,9

Ctip = 4,4 4,9 6,1

CSOB = 16,2 - -

bH/bw = - 0,309 -

Wing Sw = 950 sq-ft

(t/c)avg = 0,101 n.d. Drag Div AOK

Λχ/4 = 35,00 deg Drag Div AOK

ARw = 8,00 n.d. AOK - no flutter

Ult. Load Factor Nult = 3,5 g's

TRw = 0,25 n.d.

Winglet Height, hwglt = 6,0 ft AOK - no flutter

SplanformWinglet = 35,3 sq-ft total both panels

Airfoil Type = SCDTE "CONV", "SC", "SCDTE"

Φλαπ Σπαν ηΟ/Β φλαπ = 0,75

fraction of semispan

FlapType = D "N"one,"S"ingle,"D"ouble

Leading Edge Slat = No "Yes". "No"

Empennage Horizontal Vertical

Tail Volume Ratio, VH = 0,707 0,04

1 n.d.

Aspect Ratio, ARH = 3,50 1,15 n.d.

Tail Sweep, LH = 44,0 44,0 deg

Taper Ratio, TRH = 0,47 0,47 n.d.

Tail Length, LH = 39,5 32,9 ft

Elevator / Rudder Types = Simple SHR See comments

SASoff SM% / Tail Pos'n = 9 T See comments

D

Performance Actual Limit Crz SFC = 0,633743 - lb/hr/lb

TOGW = 84 326 - lbs

MLW = 73 652 - lbs

OEW = 38 195 - lbs

Fuel Burn = 35 566 - lbs

TOFL = 10 368 11 000 ft @TOGW

LFL = 4 989 5 000 ft @ MLW

Buffet Limited CL@ICA = 0,43 0,67 n.d.

ROC@ICA = 301 300 fpm

Wingspan = 87 120 ft

VLdg@MLW = 136 150 @1.3Vsmin

L/Dcrz = 16,77 - n.d.

DOC/pax-nmi = 0,19 $/pax-nmi

Performance Summary TOGW/Sw = 88,8 psf

TOGW/PaxNmi = 0,67 lbs/Pax-nmi

20

Thrust/TOGW = 0,35 n.d.

FB/PaxNmi= 0,282 lbs/Pax-nmi

Floor Space per Pax = #N/A sq-ft

Cargo Bay Volume per Pax = 6,71 cu-ft under-floor cargo

OEW/TOGW = 45,3% n.d.

FB/TOGW = 42,2% n.d.

PAY/TOGW = 7,7% n.d.

Cruise Conditions

Max Allowable CL@ICA = 0,67 n.d. Buffet Limited CLcrz

=CLbuffet/1.3

qCrz = 202,9 psf

VtrueCrz = 851,789 fps

Mission Requirements

Mixed-Class Pax Count = 18 n.d.

(Pax) Customer Reqm't

Still Air Range = 7 000 nmi. Customer Reqm't

Mcrz = 0,88 n.d. Design Variable

Mission Revenue Cargo = 2 630 lbs Customer Reqm't

Max Payload @ MLW = 35 000 lbs Customer Reqm't

%Mission Fuel @ MLW = 0,1 % Customer Reqm't

User Selected ICA = 41 000 ft Design Variable

Min ROC @ICA = 300 fpm Customer Reqm't

Max TOFL @ MTOGW = 11 000 ft Customer Reqm't

Max LFL @ MLW = 5 000 ft Customer Reqm't

Max VLdg @MLW = 135 KEAS Customer Reqm't

TOFL & LFL Altitude = 0 ft Customer Reqm't

TOFL & LFL Temp = 59 °F Customer Reqm't

Max Wingspan = 99,85 ft Customer Reqm't

Number of Planes Built = 500 n.d. Customer Reqm't

Wing & Tails Wing Horiz. Vertical Winglet

Tail Volume Ratio, V = NA 0,71 0,041 NA n.d.

Axial Tail Pos'n (Xc/4-Xblkhd) = NA 180 100 NA in.

Aspect Ratio, AR = 8,00 3,50 1,15 2,04 n.d.

Sweep Angle, Lc/4 = 35,00 44,00 44,00 40 deg

Planform Break Pos'n Ybreak/(b/2) = 0,2 0,00 0,00 NA n.d.

Break Taper Ratio Cbreak/Croot = NA 1,00 1,00 NA n.d.

Tip Taper Ratio, Ctip/Croot = 0,25 0,47 0,47 0,35 n.d.

Glove Extension at SOB, DC/CSOB = 0 NA NA NA 0,5

Yehudi Extension at SOB, DC/CSOB = 0,15 NA NA NA n.d.

Flap Chord Fraction Cflap/Cw = 0,18 0,25 0,25 NA n.d.

Area, total trapezoidal, S = 950 207 104 35 sq-ft

Span, b = 87 27 11 6 ft

Wing Dihedral Angle, G = 3 0,00 NA 75 deg

Vertical Offset, DZ = -40 150,00 30,00 NA in.

Airfoil t/c @ SOB = 0,150 NA NA NA n.d.

Airfoil t/c @ Break = 0,120 0,07 0,07 NA n.d.

21

Airfoil t/c @ Tip = 0,085 NA NA NA n.d.

MAC Incidence Angle = 2 -3,00 NA NA deg

Axial Wing pos'n, Xc/4trap@C/L = 454,1984 NA NA NA in. Front Spar Pos'n, XFS/C

= 0,1 NA NA NA n.d.

Rear Spar pos'n at SOB, XRS/C = 0,75 NA NA NA n.d.

Rear Spar pos'n, Break->Tip, XRS/C = 0,7 NA NA NA n.d.

Fuselage Total Fuselage Length, Lfus = 1040 in.

Nose Fineness Ratio, Lnose/Dia = 1,9 n.d.

Tailcone Fineness Ratio, Ltail/Dia = 3,5 n.d.

Tailcone Height Ratio, Htip/Hfus = 0,2 n.d.

Tailcone Width Ratio, Wtip/Wfus = 0,05 n.d.

Nose Droop Ratio, DZ/Hfus = 0,16 n.d.

Tailcone Rise Ratio, DZ/Hfus = 0,2 n.d.

Cockpit Bulkhead pos'n, DX/Lnose = 0,65 in.

Aft Blkhd pos'n, DX/Ltailcone = 0,6 in.

Cabin FloorSpace (blkhd to blkhd) = #N/A sq-ft

Usable Floor Space per Passenger = #N/A sq-ft/pax

Nacelles Number of Nacelles = 2 1-4

Fan diameter, Dfan = 48,84282 in.

Length Ratio, L/Dfan = 3 n.d.

Dia.Ratio, Dnac/Dfan = 1,25 n.d.

Inboard Outboard

Lateral pos'n Y = 0,00 80,00 in.

(Xnozzle-XLEwing) = 0,00 -20,00 n.d.

Z rel to WingLE or floor = 0,00 35,00 n.d.

Nose Gear Z pos'n of Trunnion = -30 in.

X station of nosegear = 98,0 in.

Tire Pressure = 900 psi

Number of Tires = 2 1,2

Limits

MinWng/GW @ aft c.g. = 0,04 n.d.

1g Pitch Angle = 10,0 deg

No Gear Load Pitch Angle = 12,0 deg

1g Roll Clearance Angle = 8,0 deg

Main Gear Static Ground Line pos'n

= -90,0 in.

Z pos'n of Trunnion = -46,0 in.

22

Y pos'n of Trunnion = 75,0 in.

Strut Aft-sweep, Extended = 3,0 deg

Strut Aft-sweep, Retracted = 6,0 deg

Oleo Stroke-Out from 1g = 10,0 in.

Oleo Stroke-In from 1g = 1,0 in.

X maingear = 642,0 in.


Number of Tires per Truck = 2 1,2,4

23

B. APPENDIX

Optimized model for cruise speed Mach 0.88


Ncrew = 2 n.d.

Typ 3 class


Diameter (avg), Dfus = 8,6 ft

Dfus =6,1






n.d. | lbs/ft2





Propulsion


Neng = 2 n.d.

Eng Location = F

"W"ing, "F"uselage


Bypass Ratio (BPR) = 5 n.d. **TFAN**


24



Span, b = 86,3 25,4 11,0


Croot = 17,3 9,9 13,0

Ctip = 4,3 4,6 6,1

CSOB = 16,0 - -

bH/bw = - 0,294 -



(Pax) Customer

Reqm't

Still Air Range = 7 000 nmi.

Customer Reqm't

Mcrz = 0,88 n.d.

Design Variable

Mission Revenue Cargo = 2 630 lbs

Customer Reqm't

Max Payload @ MLW = 35 000 lbs

Customer Reqm't

%Mission Fuel @ MLW = 0,1 %

Customer Reqm't

User Selected ICA = 41 000 ft

Design Variable

Min ROC @ICA = 300 fpm

Customer Reqm't

Max TOFL @ MTOGW = 11 000 ft

Customer Reqm't

Max LFL @ MLW = 5 000 ft

Customer Reqm't

Max VLdg @MLW = 135 KEAS

Customer Reqm't

TOFL & LFL Altitude = 0 ft

Customer Reqm't

TOFL & LFL Temp = 59 °F

Customer Reqm't

Max Wingspan = 99,85 ft Customer Reqm't


Wing Sw = 930 sq-ft





TRw = 0,25 n.d.




Φλαπ Σπαν ηΟ/Β φλαπ = 0,75 fraction of semispan


25




4 n.d.







D


TOGW = 83 135 - lbs

MLW = 73 633 - lbs

OEW = 38 189 - lbs


TOFL = 9 594 11 000 ft @TOGW

LFL = 4 986 5 000 ft @ MLW





L/Dcrz = 16,72 - n.d.











Cruise Conditions


=CLbuffet/1.3

qCrz = 202,9 psf






Sweep Angle, Lc/4 = 42,00 44,00 44,00 40 deg

26

Planform Break Pos'n Ybreak/(b/2) = 0,2 0,00 0,00 NA n.d.


Tip Taper Ratio, Ctip/Croot = 0,25 0,47 0,47 0,35 n.d.

Glove Extension at SOB, DC/CSOB = 0 NA NA NA 0,5

Yehudi Extension at SOB, DC/CSOB = 0,15 NA NA NA n.d.



Span, b = 86 25 11 6 ft






MAC Incidence Angle = 2 -3,00 NA NA deg Axial Wing pos'n, Xc/4trap@C/L

= 403,3681 NA NA NA in. Front Spar Pos'n,

XFS/C = 0,1 NA NA NA n.d. Rear Spar pos'n at SOB,

XRS/C = 0,75 NA NA NA n.d. Rear Spar pos'n, Break->Tip,

XRS/C = 0,7 NA NA NA n.d.

Fuselage Total Fuselage Length,

Lfus = 1040 in.











Nacelles

Number of Nacelles = 2 1-4




Inboard Outboard


(Xnozzle-XLEwing) = 0,00 -20,00 n.d.

27


Nose Gear Z pos'n of Trunnion

= -30 in.




Limits





Main Gear Static Ground Line

pos'n = -90,0 in.










28

C. APPENDIX






Dfus =6,1






n.d. | lbs/ft2





Propulsion


Neng = 2 n.d.

Eng Location = F

"W"ing, "F"uselage





29


Span, b = 86,7 20,2 10,6


Croot = 17,3 7,8 12,5

Ctip = 4,3 3,7 5,9

CSOB = 16,1 - -

bH/bw = - 0,233 -



(Pax) Customer

Reqm't


Customer Reqm't



Customer Reqm't


Customer Reqm't


Customer Reqm't



Customer Reqm't


Customer Reqm't


Customer Reqm't


Customer Reqm't


Customer Reqm't


Customer Reqm't

Max Wingspan = 105 ft Customer Reqm't


Wing Sw = 940 sq-ft





TRw = 0,25 n.d.








30


4 n.d.







D


TOGW = 84 459 - lbs

MLW = 74 161 - lbs

OEW = 38 708 - lbs


TOFL = 9 570 11 000 ft @TOGW

LFL = 4 998 5 000 ft @ MLW





L/Dcrz = 16,28 - n.d.











Cruise Conditions


=CLbuffet/1.3

qCrz = 221,8 psf






Sweep Angle, Lc/4 = 42,00 50,00 50,00 40 deg Planform Break Pos'n Ybreak/(b/2)

= 0,25 0,00 0,00 NA n.d.


31

Tip Taper Ratio, Ctip/Croot = 0,25 0,47 0,47 0,35 n.d. Glove Extension at SOB, DC/CSOB

= 0 NA NA NA 0,5 Yehudi Extension at SOB,

DC/CSOB = 0,35 NA NA NA n.d.



Span, b = 87 20 11 6 ft







Axial Wing pos'n, Xc/4trap@C/L = 451,9088 NA NA NA in. Front Spar Pos'n,





Lfus = 1108 in.















Inboard Outboard


(Xnozzle-XLEwing) = 0,00 40,00 n.d.



32




Limits






pos'n = -90,0 in.










33

D. APPENDIX






Dfus =6,1






n.d. | lbs/ft2





Propulsion


Neng = 2 n.d.

Eng Location = F

"W"ing, "F"uselage


Bypass Ratio (BPR) = 4,5 n.d. **TFAN**



34


Span, b = 84,7 19,4 10,5


Croot = 19,4 7,5 12,5

Ctip = 4,8 3,5 5,9

CSOB = 17,9 - -

bH/bw = - 0,229 -



(Pax) Customer

Reqm't


Customer Reqm't



Customer Reqm't


Customer Reqm't


Customer Reqm't



Customer Reqm't


Customer Reqm't


Customer Reqm't


Customer Reqm't


Customer Reqm't


Customer Reqm't



Wing Sw = 1 025 sq-ft





TRw = 0,25 n.d.







35



6 n.d.







D


TOGW = 92 528 - lbs

MLW = 76 768 - lbs

OEW = 41 252 - lbs


TOFL = 10 625 11 000 ft @TOGW

LFL = 5 005 5 000 ft @ MLW





L/Dcrz = 15,44 - n.d.











Cruise Conditions


=CLbuffet/1.3

qCrz = 241,5 psf







= 0,3 0,00 0,00 NA n.d.

36






Area, total trapezoidal, S = 1 025 107 97 39 sq-ft

Span, b = 85 19 11 6 ft












Lfus = 1283 in.















Inboard Outboard




Nose Gear

37

Z pos'n of Trunnion = -30 in.




Limits






pos'n = -90,0 in.










38

E. APPENDIX






Dfus =6,1






n.d. | lbs/ft2





Propulsion


Neng = 2 n.d.

Eng Location = F

"W"ing, "F"uselage





39


Span, b = 85,2 19,3 12,6


Croot = 22,7 8,8 14,9

Ctip = 5,7 4,1 7,0

CSOB = 21,0 - -

bH/bw = - 0,227 -



(Pax) Customer

Reqm't


Customer Reqm't



Customer Reqm't


Customer Reqm't


Customer Reqm't



Customer Reqm't


Customer Reqm't


Customer Reqm't


Customer Reqm't


Customer Reqm't


Customer Reqm't



Wing Sw = 1 210 sq-ft





TRw = 0,25 n.d.








40


3 n.d.







D


TOGW = 108 806 - lbs

MLW = 82 705 - lbs

OEW = 47 068 - lbs


TOFL = 10 982 11 000 ft @TOGW

LFL = 4 962 5 000 ft @ MLW





L/Dcrz = 14,01 - n.d.











Cruise Conditions


=CLbuffet/1.3

qCrz = 256,8 psf







= 0,35 0,00 0,00 NA n.d.


41





Area, total trapezoidal, S = 1 210 125 138 46 sq-ft

Span, b = 85 19 13 6 ft












Lfus = 1316 in.

Nose Fineness Ratio, Lnose/Dia = 6 n.d.














Inboard Outboard





42




Limits






pos'n = -90,0 in.










43

F. APPENDIX

Program codes for the modification tool excel macro.

Dim JetsizerFile As String Sub Iterate_AR() ' ' Iterate_AR Macro ' ' Dim MTOGW As Double, OEW As Double, FB As Double, LD As Double, CLCZ As Double Dim SLST As Double, tcroot As Single, tcbreak As Single, tctip As Single Dim Archive As Integer Dim VarVal As Variant Dim i As Integer, j As Integer, index As String Dim Flutter As String Application.ScreenUpdating = False JetsizerFile = Cells(2, 3) ParentWin = ThisWorkbook.Name Windows(ParentWin).Activate 'Change to WKname i = Selection.Row 'Always assume user selects first variable step on execution j = Selection.Column Archive = Cells(3, 3) 'list baseline archive index to load index = "C11" Do While Cells(i, j) <> Empty VarVal = Cells(i, j) Windows(JetsizerFile).Activate 'Load baseline from archive Sheets("Archive").Select

44

Range("A69").Select Do While ActiveCell.Value <> Archive ActiveCell.Offset(59, 0).Select Loop Application.Run JetsizerFile & "!RestoreFile" 'Load in new Variable Value Sheets("Configuration").Select Range(index).Value = VarVal 'Iterate on tc Do While Cells(81, 2) <> "FAIL" And Cells(23, 3) < 0.151 Cells(23, 3) = Cells(23, 3) + 0.001 Cells(24, 3) = Cells(23, 3) * 0.85 Cells(25, 3) = Cells(23, 3) * 0.7 Loop Cells(23, 3) = Cells(23, 3) - 0.001 Cells(24, 3) = Cells(23, 3) * 0.85 Cells(25, 3) = Cells(23, 3) * 0.7 'Run for Convergence Application.Run "Convergebuttons" Application.Run "reduceslst" 'Get Weights and Parameters MTOGW = Sheets("Mission").Range("I29") OEW = Sheets("Mission").Range("I31") FB = Sheets("Mission").Range("I32") LD = Sheets("Mission").Range("I39") CLCZ = Sheets("Aero").Range("I19") SLST = Sheets("Mission").Range("D19") tcroot = Sheets("Configuration").Range("C23") tcbreak = Sheets("Configuration").Range("C24") tctip = Sheets("Configuration").Range("C25") If Sheets("Mission").Range("K8") = "FLUTTER" Then Flutter = "Flutter" End If Windows(ParentWin).Activate 'Change to WKname Cells(i + 1, j) = MTOGW Cells(i + 2, j) = OEW Cells(i + 3, j) = FB Cells(i + 4, j) = LD Cells(i + 5, j) = CLCZ Cells(i + 6, j) = SLST Cells(i + 7, j) = tcroot Cells(i + 8, j) = tcbreak Cells(i + 9, j) = tctip

45

If Flutter = "Flutter" Then Cells(i - 1, j) = Flutter End If Flutter = Empty j = j + 1 Loop 'Load from archive Windows(JetsizerFile).Activate Sheets("Archive").Select Range("A69").Select Do While ActiveCell.Value <> Archive ActiveCell.Offset(59, 0).Select Loop Application.Run JetsizerFile & "!RestoreFile" End Sub Sub ConvergeButtons() Application.Run JetsizerFile & "!ConvergeCLandTOGW" loopnum = 0 maxloop = 20 indexch = 33 Do While loopnum < maxloop And indexch <= 38 Application.Run JetsizerFile & "!ConvergeCLandTOGW" indexch = 33 Do While Cells(indexch, 14) = "OK" indexch = indexch + 1 Loop Select Case indexch Case 33 Application.Run JetsizerFile & "!SLSTforTOFL" Case 34 Application.Run JetsizerFile & "!SizeSwForLFL" Case 35

46

Application.Run JetsizerFile & "!SizeSwForCLICA" Case 36 Application.Run JetsizerFile & "!SizeSLSTforROC" Case 38 Application.Run JetsizerFile & "!SizeSwForVldg" End Select loopnum = loopnum + 1 Loop End Sub Sub reduceSLST() Windows(JetsizerFile).Activate Sheets("Mission").Range("D19").Select SLSTtemp = ActiveCell.Value indexch = 39 Do While indexch > 38 SLSTnew = SLSTtemp - 1000 Sheets("Mission").Range("D19").Select ActiveCell.Value = SLSTnew indexch = 33 Do While Cells(indexch, 14) = "OK" indexch = indexch + 1 Loop Application.Run "ConvergeButtons" SLSTtemp = SLSTnew Loop End Sub Sub Iteration_Sweep() ' ' Iteration_Sweep Macro '

47

Dim MTOGW As Double, OEW As Double, FB As Double, LD As Double, CLCZ As Double Dim SLST As Double, tcroot As Single, tcbreak As Single, tctip As Single Dim Archive As Integer Dim VarVal As Variant Dim i As Integer, j As Integer, index As String Dim Flutter As String, MDD As String Application.ScreenUpdating = False JetsizerFile = Cells(2, 3) ParentWin = ThisWorkbook.Name Windows(ParentWin).Activate 'Change to WKname i = Selection.Row 'Always assume user selects first variable step on execution j = Selection.Column Archive = Cells(3, 3) 'list baseline archive index to load index = "C12" Do While Cells(i, j) <> Empty VarVal = Cells(i, j) Windows(JetsizerFile).Activate 'Load baseline from archive Sheets("Archive").Select Range("A69").Select Do While ActiveCell.Value <> Archive ActiveCell.Offset(59, 0).Select Loop Application.Run JetsizerFile & "!RestoreFile" 'Load in new Variable Value Sheets("Configuration").Select Range(index).Value = VarVal 'Iterate on tc Do While Cells(81, 2) <> "FAIL" And Cells(23, 3) < 0.151 Cells(23, 3) = Cells(23, 3) + 0.001 Cells(24, 3) = Cells(23, 3) * 0.85 Cells(25, 3) = Cells(23, 3) * 0.7 Loop

48

Cells(23, 3) = Cells(23, 3) - 0.001 Cells(24, 3) = Cells(23, 3) * 0.85 Cells(25, 3) = Cells(23, 3) * 0.7 'Run for Convergence Application.Run "Convergebuttons" Application.Run "reduceslst" 'Get Weights and Parameters MTOGW = Sheets("Mission").Range("I29") OEW = Sheets("Mission").Range("I31") FB = Sheets("Mission").Range("I32") LD = Sheets("Mission").Range("I39") CLCZ = Sheets("Aero").Range("I19") SLST = Sheets("Mission").Range("D19") tcroot = Sheets("Configuration").Range("C23") tcbreak = Sheets("Configuration").Range("C24") tctip = Sheets("Configuration").Range("C25") If Sheets("Mission").Range("K8") = "FLUTTER" Then Flutter = "Flutter" End If If Sheets("Mission").Range("K7") = "FAIL" Then MDD = "MDD" End If Windows(ParentWin).Activate 'Change to WKname Cells(i + 1, j) = MTOGW Cells(i + 2, j) = OEW Cells(i + 3, j) = FB Cells(i + 4, j) = LD Cells(i + 5, j) = CLCZ Cells(i + 6, j) = SLST Cells(i + 7, j) = tcroot Cells(i + 8, j) = tcbreak Cells(i + 9, j) = tctip If Flutter = "Flutter" Then Cells(i - 1, j) = Flutter End If If MDD = "MDD" Then Cells(i - 1, j) = MDD End If Flutter = Empty MDD = Empty j = j + 1

49

Loop 'Load from archive Windows(JetsizerFile).Activate Sheets("Archive").Select Range("A69").Select Do While ActiveCell.Value <> Archive ActiveCell.Offset(59, 0).Select Loop Application.Run JetsizerFile & "!RestoreFile" End Sub Sub Iteration_Winglet_Span() ' ' Iteration_Winglet_Span Macro ' Dim MTOGW As Double, OEW As Double, FB As Double, LD As Double, CLCZ As Double Dim SLST As Double, tcroot As Single, tcbreak As Single, tctip As Single Dim Archive As Integer Dim VarVal As Variant Dim i As Integer, j As Integer, index As String Application.ScreenUpdating = False JetsizerFile = Cells(2, 3) ParentWin = ThisWorkbook.Name Windows(ParentWin).Activate 'Change to WKname i = Selection.Row 'Always assume user selects first variable step on execution j = Selection.Column Archive = Cells(3, 3) 'list baseline archive index to load index = "F20" Do While Cells(i, j) <> Empty

50

VarVal = Cells(i, j) Windows(JetsizerFile).Activate 'Load baseline from archive Sheets("Archive").Select Range("A69").Select Do While ActiveCell.Value <> Archive ActiveCell.Offset(59, 0).Select Loop Application.Run JetsizerFile & "!RestoreFile" 'Load in new Variable Value Sheets("Configuration").Select Range(index).Value = VarVal 'Iterate on tc Do While Cells(81, 2) <> "FAIL" And Cells(23, 3) < 0.151 Cells(23, 3) = Cells(23, 3) + 0.001 Cells(24, 3) = Cells(23, 3) * 0.85 Cells(25, 3) = Cells(23, 3) * 0.7 Loop Cells(23, 3) = Cells(23, 3) - 0.001 Cells(24, 3) = Cells(23, 3) * 0.85 Cells(25, 3) = Cells(23, 3) * 0.7 'Run for Convergence Application.Run "Convergebuttons" Application.Run "reduceslst" 'Get Weights and Parameters MTOGW = Sheets("Mission").Range("I29") OEW = Sheets("Mission").Range("I31") FB = Sheets("Mission").Range("I32") LD = Sheets("Mission").Range("I39") CLCZ = Sheets("Aero").Range("I19") SLST = Sheets("Mission").Range("D19") tcroot = Sheets("Configuration").Range("C23") tcbreak = Sheets("Configuration").Range("C24") tctip = Sheets("Configuration").Range("C25") Windows(ParentWin).Activate 'Change to WKname Cells(i + 1, j) = MTOGW Cells(i + 2, j) = OEW Cells(i + 3, j) = FB Cells(i + 4, j) = LD Cells(i + 5, j) = CLCZ Cells(i + 6, j) = SLST Cells(i + 7, j) = tcroot

51

Cells(i + 8, j) = tcbreak Cells(i + 9, j) = tctip j = j + 1 Loop 'Load from archive Windows(JetsizerFile).Activate Sheets("Archive").Select Range("A69").Select Do While ActiveCell.Value <> Archive ActiveCell.Offset(59, 0).Select Loop Application.Run JetsizerFile & "!RestoreFile" End Sub Sub Iteration_Fuselage_length() ' ' Iteration_Fuselage_length Macro ' Dim MTOGW As Double, OEW As Double, FB As Double, LD As Double, CLCZ As Double Dim SLST As Double, tcroot As Single, tcbreak As Single, tctip As Single Dim Archive As Integer Dim VarVal As Variant Dim i As Integer, j As Integer, index As String Application.ScreenUpdating = False JetsizerFile = Cells(2, 3) ParentWin = ThisWorkbook.Name Windows(ParentWin).Activate 'Change to WKname i = Selection.Row 'Always assume user selects first variable step on execution j = Selection.Column Archive = Cells(3, 3) 'list baseline archive index to load

52

index = "C34" Do While Cells(i, j) <> Empty VarVal = Cells(i, j) Windows(JetsizerFile).Activate 'Load baseline from archive Sheets("Archive").Select Range("A69").Select Do While ActiveCell.Value <> Archive ActiveCell.Offset(59, 0).Select Loop Application.Run JetsizerFile & "!RestoreFile" 'Load in new Variable Value Sheets("Configuration").Select Range(index).Value = VarVal 'Iterate on tc Do While Cells(81, 2) <> "FAIL" And Cells(23, 3) < 0.151 Cells(23, 3) = Cells(23, 3) + 0.001 Cells(24, 3) = Cells(23, 3) * 0.85 Cells(25, 3) = Cells(23, 3) * 0.7 Loop Cells(23, 3) = Cells(23, 3) - 0.001 Cells(24, 3) = Cells(23, 3) * 0.85 Cells(25, 3) = Cells(23, 3) * 0.7 'Run for Convergence Application.Run "Convergebuttons" Application.Run "reduceslst" 'Get Weights and Parameters MTOGW = Sheets("Mission").Range("I29") OEW = Sheets("Mission").Range("I31") FB = Sheets("Mission").Range("I32") LD = Sheets("Mission").Range("I39") CLCZ = Sheets("Aero").Range("I19") SLST = Sheets("Mission").Range("D19") tcroot = Sheets("Configuration").Range("C23") tcbreak = Sheets("Configuration").Range("C24") tctip = Sheets("Configuration").Range("C25") Windows(ParentWin).Activate 'Change to WKname Cells(i + 1, j) = MTOGW Cells(i + 2, j) = OEW Cells(i + 3, j) = FB

53

Cells(i + 4, j) = LD Cells(i + 5, j) = CLCZ Cells(i + 6, j) = SLST Cells(i + 7, j) = tcroot Cells(i + 8, j) = tcbreak Cells(i + 9, j) = tctip j = j + 1 Loop 'Load from archive Windows(JetsizerFile).Activate Sheets("Archive").Select Range("A69").Select Do While ActiveCell.Value <> Archive ActiveCell.Offset(59, 0).Select Loop Application.Run JetsizerFile & "!RestoreFile" End Sub Sub Iteration_Bypass_Ratio() ' ' Iteration_Bypass_Ratio Macro ' Dim MTOGW As Double, OEW As Double, FB As Double, LD As Double, CLCZ As Double Dim SLST As Double, tcroot As Single, tcbreak As Single, tctip As Single Dim Archive As Integer Dim VarVal As Variant Dim i As Integer, j As Integer, index As String Application.ScreenUpdating = False JetsizerFile = Cells(2, 3) ParentWin = ThisWorkbook.Name Windows(ParentWin).Activate 'Change to WKname

54

i = Selection.Row 'Always assume user selects first variable step on execution j = Selection.Column Archive = Cells(3, 3) 'list baseline archive index to load index = "D23" Do While Cells(i, j) <> Empty VarVal = Cells(i, j) Windows(JetsizerFile).Activate 'Load baseline from archive Sheets("Archive").Select Range("A69").Select Do While ActiveCell.Value <> Archive ActiveCell.Offset(59, 0).Select Loop Application.Run JetsizerFile & "!RestoreFile" 'Load in new Variable Value Sheets("Mission").Select Range(index).Value = VarVal 'Iterate on tc Sheets("Configuration").Select Do While Cells(81, 2) <> "FAIL" And Cells(23, 3) < 0.151 Cells(23, 3) = Cells(23, 3) + 0.001 Cells(24, 3) = Cells(23, 3) * 0.85 Cells(25, 3) = Cells(23, 3) * 0.7 Loop Cells(23, 3) = Cells(23, 3) - 0.001 Cells(24, 3) = Cells(23, 3) * 0.85 Cells(25, 3) = Cells(23, 3) * 0.7 'Run for Convergence Application.Run "Convergebuttons" Application.Run "reduceslst" 'Get Weights and Parameters MTOGW = Sheets("Mission").Range("I29") OEW = Sheets("Mission").Range("I31") FB = Sheets("Mission").Range("I32") LD = Sheets("Mission").Range("I39") CLCZ = Sheets("Aero").Range("I19") SLST = Sheets("Mission").Range("D19") tcroot = Sheets("Configuration").Range("C23")

55

tcbreak = Sheets("Configuration").Range("C24") tctip = Sheets("Configuration").Range("C25") Windows(ParentWin).Activate 'Change to WKname Cells(i + 1, j) = MTOGW Cells(i + 2, j) = OEW Cells(i + 3, j) = FB Cells(i + 4, j) = LD Cells(i + 5, j) = CLCZ Cells(i + 6, j) = SLST Cells(i + 7, j) = tcroot Cells(i + 8, j) = tcbreak Cells(i + 9, j) = tctip j = j + 1 Loop 'Load from archive Windows(JetsizerFile).Activate Sheets("Archive").Select Range("A69").Select Do While ActiveCell.Value <> Archive ActiveCell.Offset(59, 0).Select Loop Application.Run JetsizerFile & "!RestoreFile" End Sub Sub Iteration_SLST() ' ' Iteration_SLST Macro ' Dim MTOGW As Double, OEW As Double, FB As Double, LD As Double, CLCZ As Double Dim SLST As Double, tcroot As Single, tcbreak As Single, tctip As Single Dim Archive As Integer Dim VarVal As Variant Dim i As Integer, j As Integer, index As String Application.ScreenUpdating = False JetsizerFile = Cells(2, 3)

56

ParentWin = ThisWorkbook.Name Windows(ParentWin).Activate 'Change to WKname i = Selection.Row 'Always assume user selects first variable step on execution j = Selection.Column Archive = Cells(3, 3) 'list baseline archive index to load index = "D19" Do While Cells(i, j) <> Empty VarVal = Cells(i, j) Windows(JetsizerFile).Activate 'Load baseline from archive Sheets("Archive").Select Range("A69").Select Do While ActiveCell.Value <> Archive ActiveCell.Offset(59, 0).Select Loop Application.Run JetsizerFile & "!RestoreFile" 'Load in new Variable Value Sheets("Mission").Select Range(index).Value = VarVal 'Iterate on tc Sheets("Configuration").Select Do While Cells(81, 2) <> "FAIL" And Cells(23, 3) < 0.151 Cells(23, 3) = Cells(23, 3) + 0.001 Cells(24, 3) = Cells(23, 3) * 0.85 Cells(25, 3) = Cells(23, 3) * 0.7 Loop Cells(23, 3) = Cells(23, 3) - 0.001 Cells(24, 3) = Cells(23, 3) * 0.85 Cells(25, 3) = Cells(23, 3) * 0.7 'Run for Convergence Application.Run "Convergebuttons" 'Get Weights and Parameters MTOGW = Sheets("Mission").Range("I29") OEW = Sheets("Mission").Range("I31")

57

FB = Sheets("Mission").Range("I32") LD = Sheets("Mission").Range("I39") CLCZ = Sheets("Aero").Range("I19") SLST = Sheets("Mission").Range("D19") tcroot = Sheets("Configuration").Range("C23") tcbreak = Sheets("Configuration").Range("C24") tctip = Sheets("Configuration").Range("C25") Windows(ParentWin).Activate 'Change to WKname Cells(i + 1, j) = MTOGW Cells(i + 2, j) = OEW Cells(i + 3, j) = FB Cells(i + 4, j) = LD Cells(i + 5, j) = CLCZ Cells(i + 6, j) = SLST Cells(i + 7, j) = tcroot Cells(i + 8, j) = tcbreak Cells(i + 9, j) = tctip j = j + 1 Loop 'Load from archive Windows(JetsizerFile).Activate Sheets("Archive").Select Range("A69").Select Do While ActiveCell.Value <> Archive ActiveCell.Offset(59, 0).Select Loop Application.Run JetsizerFile & "!RestoreFile" End Sub

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