Stevens2007 - Stevens Institute of Technology · · 2007-03-24The needs vs. metrics matrix ......

Stevens Institute of Technology

2007 SEA Aero Design East Report i

Stevens2007

Team Number: 053


Team Members: James Koryan, Joseph Lojek, Justin Sommer, Ramy Ghaly

Stevens Institute of Technology Stevens2007 – Team # 053

iii

Abstract

This report presents a design for an aircraft capable of lifting the maximum load; maximizing the

payload ratio. The design was developed according to the specifications of the 2007 Aero Design

competition rules. An optimal design was developed by maximizing the aerodynamic lift and reducing

the aircraft weight without compromising performance. An RC controlled plane is currently being

constructed for participation in the competition.


iv

Table of Contents

Compliance ii Abstract iii Introduction 1 Design Approach 2 Design 4 Thought Process 4 Wing Design 4 Wing Shape 5 Airfoil Selection 5 Control Surfaces 6 Wing Strategy 7 Fuselage 7 Landing Gear 8 Engine Cowl 8 Tail-plane 8 Construction 9 Wing 9 Fuselage 10 Construction Strategy 10 Tail Construction 11 Landing Gear Construction 11 Electronics used & Wiring Needed 11 Final Construction Touches 12 Calculations 12 Stability 17 Payload Prediction 19 Conclusion 19 Lessons Learned 19 Additional Sources 20 Software Utilized 20 References 20 Literature 20 World Wide Web 20 Appendix A - Payload Prediction Plot 21 Appendix B –Electronic Plan 22 Appendix C – EES Payload Prediction Calculations 23


2007 SEA Aero Design East Report 1

INTRODUCTION

Research and design began shortly after the start of the 2006-2007 school year, immediately after the

team was assembled. The team consists of four members each focusing in mechanical engineering at

Stevens Institute of Technology. The project has been the teams focus for the senior design project and

all requirements were to be completed by the team. Many hindering factors were faced but were

overcome with dedication and perseverance. The team noted that scheduling would be one of the

hardest obstacles since each member had drastically varying schedules. Tools such as Gantt charts

were applied to the process to help with time management and guidance to a successful project. Project

management and organizational skills were essential to the success of the project.

With time and responsibilities allotted to each member further work commenced. Brainstorming,

research and calculations soon became the next step of focus. Various sources were used in educating

the team in fluid mechanics, aerospace design, and the construction of an airplane. Research involved

reading and studying an assortment of texts and publications in the design of RC model planes and flight

theory. Drafting ideas and modeling airfoil types quickly followed the team’s research.

Using a multitude of software and mathematical models, optimization of the final design began. Once

the optimum design was obtained construction of the fuselage, wing, and other components followed.

Using simple craftsmen techniques, tedious labor, and extreme attention to detail, the design team

slowly assembled the airplane. The final components were then mono-coated and the plane was

prepared for test flights.



DESIGN APPROACH

To focus the team’s initial ideas into a weighted concern, a list of needs and metrics was created. The

list of needs consisted of all the necessary concerns involved in designing a compliant aircraft to this

year’s competition and successful design. The list of needs can be seen below in Table 1.

Table A - Needs Table

Following the construction of the needs, the team instituted a metrics of critical properties of the aircraft to be weighted as noted Below in Table B.

Table B - Needs-Metrics Table



The next step was to weigh the needs versus metrics and then observe the focal points of the teams

design. Each need and metric was assigned a value according to its importance. The needs and metrics

that had the most importance received the higher values. The needs vs. metrics matrix (Table C)

provided the team with a clear view of the most important factors of the design.

Table C – Need vs. Metrics Matrix



DESIGN

Thought Process

Soon after receiving this assignment the team sought out reference materials in radio controlled model

airplane design and construction as well as advice from several professors. In previous Aero Design

competitions attended by Stevens Institute of Technology several entries have been overly designed and

didn’t perform as expected. The team gathered that the problem behind this event was concerning the

fabrication of a plane that accurately resembles the original concept. The accuracy to which we construct

our plane will govern the predictability of its performance. The main goal of the team was to design a

plane that carries the maximum weight possible. While taking under consideration the problems that

may arise during the construction phase, the team decided to implement the simplest most effective

designs and concepts.

Wing Design

There are several types of airplane configurations that fit with in the specifications of the competition

that the team had to choose from. The team spent time brainstorming and presented ideas to each other

which consisted of wing configurations, airfoil profiles and other body configurations. The team’s first

decision was between the monoplane, bi-plane, tri-plane, and delta wing designs. The team chose the

monoplane design because it was the most common design and the design that the team presumed to be

easiest to construct.



Wing Shape

The team also decided to implement a straight wing shape as opposed to the elliptical, tapered, or swept

wing shapes. Although the elliptical, tapered, and swept wing shapes provide better spanwise lift

distribution and a reduction in wing tip losses the team opted for the straight wing design because of its

ease of construction. However to compensate for the differences between the straight wing shape and the

other shapes the team researched several methods of reducing wing tip losses and optimizing spanwise

lift distribution. Winglets, endplates and Hoerner tips were among the methods discussed by the team.

As stated before the team chose the simplest and most effective design which was the endplates. The

winglet and the Hoerner tip provide similar effects as the endplate but the group decided that the

complicated design would not be feasible to construct in the allotted time. From here the team

researched airfoils that were designed to have high lift coefficients.

Airfoil Selection

The team used programs such as Xfoil and WinFoil as well as internet resources to analyze and acquire

the best airfoil for our design. The team also analyzed the airfoils used in previous competitions. We

gathered that an airfoil with a slender trailing edge would pose difficulty in reproducing with the current

equipment available to us. The Eppler 423, the Selig 1210, the Aquila and the Clark Y (Figure 1) were

among our top choices. The Eppler 423 and the Selig 1210 have very high coefficients of lift, unlike the

Aquila and Clark Y profiles. However the Aquila and the Clark Y have simpler geometries because of

the thicker trailing edge and lower camber, making them the easiest to build accurately. The team chose

the Eppler 423 because it has the highest lift coefficient and a thicker trailing edge than the Selig 1210

(Figure 2).



Figure 1 - Lift Coefficient vs. Angle of Attack

Figure 2 – Eppler 423, Selig 1210

Control Surfaces

The team believed that the proper construction of the control surfaces will have the greatest effect in the

performance of the plane. The flaps, ailerons, elevator and the rudder control surfaces have been

designed to be simple and effective. The team used proportions similar to the suggested proportions for

flaps and ailerons from the Basics of R/C Model Aircraft Design by Andy Lennon. The team referenced

this book in several other applications to be mentioned later in the text. Calculations were also done to

Cl Vs. AOA

-0.5

0

0.5

1

1.5

2

2.5

-6 -4 -2 0 2 4 6 8 10 12 14 16

AOA

Cl

E423 Re: 200000 AQUILA, Re:101100 AQUILA, Re:150500 AQUILA, Re:203900 AQUILA, Re:301100 Clark Y, Re:63100

Clark Y, Re:20380 E-423, RE: 60300 E-423, RE: 198600 E-423, RE: 296900



ensure that the control surfaces were effective and that the plane would fly. The calculations will be

further explained later in the text as well. There were several types of flaps that were discussed as well

as leading edge heavy lift devices. Implementing a complicated flap design would prove beneficial with

respect to the increased lift and other flight performance characteristics. Nonetheless the team decided to

implement a plain-flap design because of its simplicity. The leading edge heavy lift devices were taken

out of consideration due to our lack of knowledge and the technical skills to construct them properly.

Wing strategy

The full wing span is 80 inches and the rectangular fuselage has a width of 6 inches. Since the design of

the shape of the body has been chosen to be a high-wing monoplane, the team members have considered

the following methods in constructing the wings relative to the fuselage The first method was to

construct the whole wing then have it glued to the top of the fuselage. The next method was to cut out a

groove in the fuselage for mounting the wing. The groove is shaped according to the airfoil dimensions

in efforts to reduce wing body losses. The last method consisted of a three part wing which comprised of

a left, right, and middle section. The middle section is permanently attached to the fuselage and the other

two wing sections will slide into position and locked in place using pins and slots. This enables the team

to detach the wings for easy transportation and more importantly the capability of replacing damaged

wing sections. Also, the middle section will be properly reinforced to handle the wing loading.

Fuselage

The fuselage has a rectangular cross section which tapers off after the payload area. Originally the team

considered a circular or elliptical cross-section that tapers off towards the tail because it has less drag

than a rectangular cross-section fuselage. The team ran into complications determining how to attach the

landing gears and the middle wing section to the fuselage with out creating too much drag and

compromising hull integrity. The landing gear needed to be mounted to a flat surface. The team



contemplated mounting the landing gears to a rectangular frame located inside the fuselage and then

cutting holes in the exterior shell to allow the landing gear to protrude out. The holes obviously will

increase the drag and the deformation of the landing gears during landing may break the exterior shell.

The rectangular cross-section fuselage provided a flat surface required to mount the landing gear without

an added difficulty. The wing construction and assembly will be discussed later in the text. The engine

will be mounted to a section of ply-wood and an

Landing Gear

The tri-cycle landing gear layout was chosen over the tail-dragger layout because it has superior steering

during take off and landing and it’s the most commonly used layout. The team considered different

types

Engine Cowl

The engine cowl will cover the engine to reduce drag created by the frontal area of the fuselage. The

team speculated that the best engine cowl would one provide a smooth transition from the nose to the

fuselage. A conic shaped engine cowl would have been ideal if then team had continued with a

cylindrical shaped fuselage. The team decided to buy the cowl because. In addition, the price of the

engine cowl is relatively cheaper than the price the student would have had to pay for time and effort put

in constructing it.

Tail-plane

The team will be using a “T-Tail” design for the final designed aircraft. It was debated whether or not to

use other designed tails such as a “V-Tail” or a tail with its horizontal wing below the vertical rudder.

All types were analyzed and studied for their theoretical traits. It was the “T-Tail” design that proved to

be most beneficial for our application.



CONSTRUCTION

After the team chose which of the conceptual designs to apply to the final design, the members started

the construction process. First, the group members thought of different methods to construct the

airplane’s components as well as how to assemble all of them together. Below is a discussion of the

different methods thought of by the group members and the optimal method that was chosen.

Wing

The team began with making stencils of the Eppler 423 airfoil profile. The stencils were cut from 1/8"

balsa wood sheets and formed the ribs. The left and right wing sections have twelve ribs that span a

total length of 33”. The ribs were mounted on carbon fiber tubing; as described in the design section in

this report. The flaps and ailerons with a cord length of 2.75 inches will span the majority of the wing.

A 1/2" hole was then made at the theoretical mean aerodynamic point of the airfoil where a 3/8" carbon

fiber tube was then inserted and glued. It is at the mean aerodynamic point where all forces are assumed

to be acting. A second hole was then drilled in the ribs. The ribs were fixed in their proper places. Two

carbon fiber tubes were put through the holes. Before gluing the ribs to the tube, the ribs were first fixed

in a clamping system designed by the group members to ensure there stability for a day until the glue

hardens.

As of this date, March 21, 2007, the team has only finished the above constructions. However, in the

following weeks the team is aiming to finish constructing the wings. The two flap portions on each side

of the wings will be connected together through the fuselage with a carbon fiber tube that runs through

the holes made in the ribs of the flaps. Holes were also made in the ailerons’ ribs where each aileron on

each side was made by connecting the ribs by carbon fiber tubes. The reason two separate tubes were

used is because each aileron will be connected to a different servo since they will be deflected in



different direction with different angles. Thin balsa sheets well by used to cover the skeletons of the

fixed portion of the wing, the flaps, and the ailerons.

Fuselage

This section contains the teams’ considered construction methods and will list and justify the reasons

why the chosen method is optimal. The team considered constructing rectangular ribs and connecting

them with four carbon fiber tubes that run through each corner of the ribs and building a frame and

cover it with thin sheets of balsa.

The previous considerations were listed in order to compare to the chosen method for justification

purposes. The team is assuming that the airplane will drop a distance of 3 feet while landing. The impact

that is exerted on the landing gear will be delivered to the bottom side of the fuselage. It is known from

previous designs that balsa wood will not be the best suited material to use to absorb this impact force

that is exerted upon landing. As a result the team members have decided to use a stronger material to

absorb the impact. The team decided on a plywood due to its relative light weight and strength. It is also

one of the easiest materials to work with; easy to cut and glue.

Construction strategy:

As of this date, the team did not yet start construction of the fuselage. However, the strategy that the

team is to follow is a: 6”x48” rectangles will be cut on 1/4" balsa sheets. The three sides will be glued at

their corners with the use of an L shape light balsa sticks. The bottom side of the fuselage will be made

out of plywood. As described above, 4 inches of each wing on either side will be fixed to the fuselage.

They will be glued and reinforced by using carbon fiber tubes that run through the fuselage and the wing

set on both sides of the fuselage.



Tail construction:

The method of construction to be used for the horizontal tail is similar to that of the main wing. The

chosen design and the strategy to be followed are similar to the main wing. The ribs will be a profile of

NACA 0009 airfoils and constructed from balsawood. Fiber rods will once again be used to support the

construction of the tail and finally thin, light balsa sheets will be used to cover the skeleton. The same

will be done to the rest of the horizontal tail; however, it will be fixed relative to the body of the

airplane.

The vertical tail will be made out of a vertical NACA0009. At its top, the horizontal tail will be fixed. Its

bottom will be fixed to the top of the fuselage by strong glue and reinforced with L shape pieces of

stronger balsa.

Landing gear construction:

The school, Stevens Institute of Technology, currently has two sets of commercial landing gears and 4"

wheels that were used in previous years and have showed an accepted performance. As a result, the team

has decided to use them to save time, effort, and cost. In addition, the available landing gears are made

with fine finish to decrease the surface drag induced.

Electronics used and wiring needed:

The team will be using five servos to control the airplane’s motion on ground as well as in midair

through the usage of a remote control. One servo will be controlling the two flaps. Two separate servos

will be used to control the ailerons since the two ailerons will be deflected in different directions as well

as with different magnitudes. Another servo will be connected to both the rudder and the front wheel.

The last servo will be connected to the horizontal tail. All the wires will that connects the servos to the

airplanes components, are hidden inside the fuselage and hidden for scenic purposes.



Final construction touches:

Finally, the team will mono-coated all members to obtain the finest surface finish to minimize surface

drag. In addition, mono-coating will increase the rigidity of the relative parts of the airplane since it will

hold the relative parts together.

Calculations

The group utilized as many resources as possible for the plane's theoretical calculations, including, but not limited

to, "Basics of R/C Model Aircraft Design" by Andy Lennon, the "white paper" written by Dr. Leland M. Nicolai,

"Fluid Mechanics" by Frank White, and the SAE website. Previous groups from SIT predominantly used the

"white paper" as a general guideline; under the advice of our advisor, Professor Thangam, the team used the

guidelines presented in Lennon's R/C Model book and mathematical models from the “Fluid Mechanics” textbook

in addition to Nicolai’s “white paper”. While the "white paper" produces a sufficient model, the team agreed that

it would be more accurate and practical to use empirical data for smaller model planes (provided in Lennon's

book) where possible rather than to try and use formulas associated with larger aircraft. Aircraft calculations do

not scale down accurately, and the group has adjusted our model to reflect this, replacing some calculations with

experimental data.

The group utilized EES to assemble the mathematical models; these models have been adjusted and refined many

times in order to produce the most accurate performance prediction possible. The models have now been

arranged so that the input entries are the runway distance, the weight, air density constants, and dimensions of the

aircraft; from this, EES determines takeoff velocity, the thrust required to takeoff at that velocity, the engine

thrust at that velocity, and also the runway distance required for landing.

The group uses two basic models to calculate performance. The first comes from “Fluid Mechanics”. The group

used a force balance in the direction of takeoff which is



RkVTRollingDragThrustdtdvmFs −−=−−== 2 [EQ-2]

where R is the rolling resistance, and k is defined as follows:

∑= effectiveD ACk ρ21

. [EQ-3]

The second model comes from Nicolai’s “white paper”, and it uses the mean acceleration for takeoff instead of

dV/dt. This turns our force balance into

)( rolldmean FFTma −−= , [EQ-4]

where mean

TO

aV

s2

2

= , [EQ-5]

and s is the distance allowed for takeoff.

Each aspect of the plane (the wings, fuselage, landing gear, flaps, etc.) adds to the total drag force and must be

accounted for in order to depict our plane accurately. The CD values for each respective aspect were calculated

using the model in Nicolai’s “white paper” and the results are tabulated in Table E.

CD values

Table E – Drag Coefficients



In order for this model to produce an accurate prediction, the drag coefficients and the effective areas must be as

close as possible to the actual areas during taking off. During takeoff and landing, the effective lifting area is

reduced because the flaps are down (which must be incorporated into the stall velocity calculation), and more

blockage drag must also be taken into account (distance calculation). The ΣCDAeffective calculation was performed

with

)()()()()( LGdLGvwingdvwinghwingdhwingplanformdwingfusedfuseeffectived ACACACACACAC ++++=∑

)()()( flapsdownkagedflapsblocafrontalArekagedfrontblocenginedengine ACACAC +++ [EQ-6]

for performance prediction during takeoff and landing. The angle of deflection for the flaps during takeoff was

taken to be 13 degrees and during landing to be 20 degrees. This has been accounted for in terms of wing

planform area reduction and also in terms of frontal blockage (creating more drag). It also is incorporated in the

calculation of the lift coefficient. The lift coefficient is a function of the max camber of the airfoil (in our case,

the Eppler 423), the angle of attack of the wing, and the angle of deflection of the flaps. CL is calculated as

follows:

)2sin(/)2sin(*max/ ch

chCC flapLflapsLw +++= ααα [EQ-7]

The coefficients of lift and the ΣCDAeffective values for the different parts of our flight are in the following table.

Lift Coefficients

Table F – Lift Coefficients

The group decided to simplify the model for takeoff by ignoring the Rolling resistance because it is very small

and has little effect on the overall takeoff calculation. This reduces the model to



2kVTDragThrustdtdvmFs −=−== [EQ-8]

and )( dmean FTma −= [EQ-9]

where kV2 is the drag force at takeoff. The takeoff velocity is easily obtained by first obtaining the stall speed

(Vs). The stall speed is found with

21

)2(PLMax

S ACWVρ

= , [EQ-10]

where CLMax is the lift coefficient with the flaps and ailerons in takeoff position, ρ is air density at sea level, and

AP is the planform area of the wing. In order to insure takeoff, the takeoff velocity must be at least

STO VV 2.1= [EQ-11]

On another note, the SAE rules dictate a takeoff runway of only 200 feet, which means that distance is important

to determine because it’s a constraint; however, the time required to takeoff is not. Therefore, we substitute the

following into our model,

dsdVV

dtdV

= . [EQ-12]

After separating variables and integrating, we discover that

∫∫ −=

TOVS

kVTVdVmds

02

0 )(2, [EQ-13]

and 2ln2 TO

o kVTT

kmS

−= ; [EQ-14]

we can set the takeoff distance to be 190 feet (leaving 10 feet for error), and solve for the thrust required to

takeoff with mass m.



The next problem to address was to determine what thrust our engine could provide. By definition, power is

VelocityThrustPower *= ; [EQ-15]

the power output of the engine is known, and so is the velocity (VTO), and so the thrust available from the engine

can be determined. The group then compared the required thrust for certain weights and matched it to find the

maximum load our plane can lift off the ground. The results are tabulated in the following table.

Thrust

Table G - Trust

When the plane weighs a total of 25 lbs, both mathematical models produce a required thrust that is less than the

available thrust. Anything more and we surpass the payload limit. Therefore, we will assume a maximum weight

of 25 lbs for takeoff.

Next, the cruising velocity must be calculated. At level flight, the weight will be the same as the force of lift.

This translates to

pL ACVLiftlbsWeight 2

2125 ρ=== , [EQ-16]

and, after solving for velocity using a lift coefficient of 2.018, an air density for an altitude of 1000 ft above sea

level, and an unflapped wing planform area. This yields a cruise velocity of 41.91 ft/s (28.6 mph). This is a very

reasonable velocity for model aircraft.



All that is left to be calculated is the landing distance. For this calculation, the group referred to Anderson’s

“Introduction to Aerospace”. For the landing distance,

VTR

T

L LWDragg

WVS

7.0

2

)]([2

)(

−+=

μ [EQ-17]

The sum of the forces in the denominator should be the instantaneous value, but for simplification, we take the

average value (which is at 0.7VT). VT is the velocity that the plane comes in at to land. To assume a factor of

safety, we take this to be

StallT VV 7.1= [EQ-18]

While there is also still a thrust from the propeller (we cannot simply shut off the engine), we can assume that it

will simply increase our landing distance by a small factor. From our model, the group attained a landing distance

of 57.3 ft (well within the allowed limit of 400 ft). Increasing this distance by 50% (due to the ignored engine

thrust) puts the distance to 86 ft, which is acceptable.

Stability

The pitching moment plays an essential role in a stable and level. The pitching moment was determined

using the equation (EQ-19) found in Basics of R/C Model Aircraft Design by Andy Lennon. The team

analyzed the effects of speed and chord length on the pitching moment (Table H). The wing was found

to have a nose down moment and due to its high camber the moment was rather large.

Pitching Moment = (CM * s * V2 * S * C) / 3519 [EQ-19]



Table H – Sample pitching moment

Where CM is the pitching moment, s is the density of air at sea level s=1, S is the wing area in in2, and V

is the velocity in mph. The pitching moment was then used to determine the horizontal tail area (HTA)

needed. EQ-20 was from the Lennon book as well. The wing area (WA), the mean aerodynamic chord

(MAC) and the HTA were used to determine the length of the tail moment arm in EQ-3.

HTA = (2.5 * MAC * 0.20 * WA) / TMA [EQ-20]

TMA = (2.5 * MAC * 0.20 * WA) / HTA [EQ-21]

Tail Moment Arm in. Wing area: 880 in2 Wing area: 800 in2

Wing Chord Length in. HTA at 180 in2 HTA At 200 in2 10 36.67 20 11 40.33 22 12 44 24

Table I – Sample Tail Moment Arm Length

The team was able to determine that the plane will have a level flight.

Pitching moment lbs/in Wing area: 880 in2

Take – Off Cruise Chord Length

(C) in. at 20 mph at 50 mph

10 -21.61 -135.09 11 -23.78 -148.6 12 -25.6 -162.1



Payload Prediction

The Payload prediction graph (see Appendix A) was generated using an EES code (see Appendix D).

This information states the maximum allowable weight that could be lifted at increased altitudes. The

graph portrays a linear plot with a Predicted Payload = {31.02-0.001 x Density Altitude}.

Conclusion

In conclusion, the team was able to apply the skills and knowledge attained at Stevens Institute of

Technology. Through strong team leadership and dedication the team was successful in their efforts.

Project management played a key role in the completion of the design and construction. It allowed the

team to meet all deadlines and organize responsibility within the group. Also, with the aid of the team’s

advisor and other noted references the construction of the aircraft was made feasible. Based on the

construction of the plane and theoretical calculations the team feels that the plane will do very well in

this year’s competition.

Lessons Learned

The unique experience of this project enabled the team to obtain many new skills. There was a great

deal of knowledge and lessons learned involving the brainstorming process. After a semester of

working as a team we discovered that brainstorming is needed to lead a strong and competitive design.

Discussion techniques were often used when developing concepts for design and as a group we were

capable in focusing our thoughts into one collective idea. Team unity strengthened greatly over the

course of the term. We were able to obtain a wider understanding of each others working habits and

weaknesses. Knowing where the group lacked strength enabled the team to increase the concern in

those areas.



Additional Sources

Software Utilized

Engineering Equation Solver (EES), Microsoft Office, Solid works, Winfoil, Matlab, Xfoil

References

Literature

F. M. White, Fluid Mechanics, 5th edition, McGraw-Hill, New York, NY, 2003

Andy Lennon, Basics of R/C Model Aircraft Design

Nicolai, Dr. Leland M, “The White Paper”

Tower Hobbies, Tower Talk, Issue #6, December 31 2006

World Wide Web

www.sae.org

Stevens Institute of Technology Stevens 2007 Team 053A di Ay = ‐0.001x + 31.01

R² = 0.999

35.00

Stevens Institute of Technology ‐ Stevens 2007 ‐ Team 053Payload Weight vs Density Altitude

Appendix A

30.00

20.00

25.00

t [lb]

15.00

Payloa

d Weight

10.00

0.00

5.00

0.00

0 1000 2000 3000 4000 5000 6000

Density Altitude [ft] 21

Stevens2007

D

C

B

AA

B

C

D

12345678

8 7 6 5 4 3 2 1

SHEET 1 OF 1

REV

UNLESS OTHERWISE SPECIFIED:

SCALE: 1:15 WEIGHT:

Team: 053PROHIBITED.

PROPRIETARY AND CONFIDENTIAL

NEXT ASSY USED ON

APPLICATION<INSERT COMPANY NAME HERE> IS WITHOUT THE WRITTEN PERMISSION OF

DWG. NO.

REPRODUCTION IN PART OR AS A WHOLE B<INSERT COMPANY NAME HERE>. ANY

SIZEDRAWING IS THE SOLE PROPERTY OF

TITLE:

THE INFORMATION CONTAINED IN THIS


THREE PLACE DECIMAL

TOLERANCING PER:

MATERIAL

FINISH

DRAWN

CHECKED

ENG APPR.

MFG APPR.

Q.A.

COMMENTS:

DATENAME

DIMENSIONS ARE IN INCHESTOLERANCES:FRACTIONALANGULAR: MACH BEND TWO PLACE DECIMAL

INTERPRET GEOMETRIC

DO NOT SCALE DRAWING

Class Regular

Data SummaryWing Span 80"

Empty Weight 10 lbEngine

Make/Model O.S. 61FX

Appendix B - Electronic Plan

11.00

2.75

4.50

80.0

0

15.0

0

6.00

12.00

3.24

24.00

6.00

6.25

7.69

6.25

64.63

12.12

jlojek

Text Box

Page 22



Appendix C

Payload Prediction Constants" E=700 "Elevation of Fort Worth, TX" R=1715.7 "Universal Gas Constant" B=.003566 "Adiabatic Lapse Rate" g=32.2 "Force of Gravity" T_0=518.69 "Temperature at Sea Level, Rankine" p_a=2116.2 "Atmospheric Pressure at Sea Level" mew=3.7373*10^(-7) "Air Viscosity" c=11 "Chord Length, Inches" MSL=1000 "Mean Sea Level, Feet" A_p=880/144 "Planform Area, Feet^2" V=52.93 "Cruise Velocity" C_L=2.018 "Lift Coefficient at Cruise" W=10 "Empty Weight of the Aircraft" "Calculations" AGL=MSL+E "Above Ground Level" T=T_0-B*MSL "Adjusted Temperature" p=p_a*(1-(B*MSL/T_0))^K "Adjusted Pressure" K=g/(R*B) "Equation Simplification" rho=p/(R*T) "Air Density" DA=145366*(1-(17.326*P/T)^.235) "Density Altitude" Re=(rho*V*c/12)/mew "Reynolds Number" q=.5*rho*(V^2) "Calculated Value" Lift=(C_L*rho*A_p*V^2)/2 "Lift Force" Payload=Lift-W "Payload Calculation"

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