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2007 SAE Aero Design East Report Group 5: Heavy Lift Cargo Plane

Stevens2007 Stevens Institute of Technology

Team Number: 053 May 2, 2007

Team Members: Joseph Lojek James Koryan Justin Sommer Ramy Ghaly

“We pledge our honor that we have abided by the Stevens Honor System”

Stevens Institute of Technology Stevens2007 – Team # 053

ii

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 has been constructed for

participation in the competition.


iii

Table of Contents

Abstract ii 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 Tail-plane 8 Construction 8 Wing 8 Wing Mounting 9 Fuselage 10 Construction Strategy 10 Tail Construction 10 Landing Gear Construction 11 Electronics used & Wiring 11 Final Construction Touches 11 Calculations 12 Stability 18 Payload Prediction 19 Prototype Testing 19 Conclusion 21 Lessons Learned 21 Additional Sources 22 Software Utilized 22 References 22 Literature 22 World Wide Web 22 Appendix A - Payload Prediction Plot 23 Appendix B –Electronic Plan 24 Appendix C – EES Payload Prediction Calculations 25 Appendix D – EES Program I: Drag and Lift Coefficients 26 Appendix E – EES Program II: The One 30


2007 SEA Aero Design East Report 1

INTRODUCTION

The main phases of the project were the research, design and construction phases. Research began

shortly after the start of fall 2006 semester, immediately after the team was assembled. Research

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

and flight theory. The team had also used the wealth of knowledge found in the Mechanical

Engineering department faculty members to formulate the team’s design approach. The team’s design

approach, in general was to design a simple and effective plane. Drafting ideas about different aspects

of the aircraft quickly followed the team’s initial research.

The design phase consisted of analysis and selection of each of the main components of the aircraft: the

wing, the airfoil, the control surfaces, the fuselage, the landing gear and the tail-plane. The team selected

a high-wing mono plane with T-tail design for various reasons stated later in the report. The Eppler 423

and the NACA 0009 airfoils were selected for the wing and horizontal tail plane respectively. 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 team began construction with the wings followed by the

fuselage and then the tail plane. Servos were installed after the majority of the plane had been

assembled. These servos were used to actuate the controls surfaces. The final components were then

mono-coated and the plane was prepared for test flight. The flight tests took place at the Moonachie

Flight Club.



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.

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

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 model, 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 mate them successfully together. Below is a discussion of the different

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

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.

Wing mounting:

After the construction of the wings were complete. A system was designed to allow proper mounting of

the wing assembly. The wing mounting assembly had to be constructed so it would situate the wings at

their proper angle of attack, which was calculated to be 13°. A pinning system was then constructed to

interlock the wings to the mounting assembly. The mounting assembly consisted of two side walls

constructed from 1/16 inch birch plywood in the shape of the Eppler 423 airfoil that was properly angled

at 13°. Then a two inch by eight inch stud was cut to its proper dimensioning in order to act as a spacer

between the two sheets of plywood. Prior to gluing the spacers to the plywood walls, holes were drilled

throughout the entire length of the material and carbon-fiber tubing was inserted. Finally all

components were secured properly with epoxy and mounted to the top of the fuselage. The wings were

then slid into the channels created in the wing mounting assembly and pinned into position at their

designed angle of attack.



Fuselage

The team assumed that the airplane will drop a distance of 3 feet while landing. The impact that will be

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 to use 1/16 inch birch 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:

The fuselage will consist of a rectangular body which will span a total distance of 60 inches from nose

to tail. The dimensions of the nose are 6 inches by 6 inches and begin to taper off at roughly 20 inches

to 3 inches by 3 inches at the tail. The primary material used in its assembly was 1/16 inch birch

plywood to increase the crafts stability and strength. Areas where the landing gear and engine are

mounted were further enforced with 3/32 inch plywood and ¼ inch plywood, respectively. This design

allowed the fuselage to withstand the stress of landing with load and solved key issues of high torsion in

the boom design from previous years.

Tail construction:

The method of construction to be used for the horizontal tail is similar to that of 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 was constructed of a stiff frame consisting of ½ inch by ½ inch wooden dowls . At its

top, the horizontal tail will be fixed. Its bottom will be fixed to the top of the fuselage by an epoxy and

reinforced with several screws to insure a solid structural bond.

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:

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.

Final construction touches:

Finally, the team mono-coated all members to obtain the finest surface finish minimizing surface drag.

The mono-coating will also make the aircraft more aesthetically pleasing, allowing it to have a fine

glossy coat. Lastly, the final preparations will be made in or to prepare the airplane for test flights.

Checks will be made to all the servos and their functionality. The engine will be once again checked

and the final touches will be conducted in order to ensure success.



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]

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



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.

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}.

Prototype Testing

The plane was first tested for balance to determine the actual location of the center of gravity. This test

was performed by first finding the balance point manually (by balancing on a finger), and then using that

point as the spot to tie a string and hang the plane from it. The center of gravity was found to be within

an inch of the aerodynamic center (AC) of the wing (between the AC and the tail), which is within the

acceptable range according to Lennon’s Basics of R/C Model Aircraft Design. This was a test while

plane was “empty”, meaning without cargo. The cargo bay is directly underneath the center of gravity,

and only reinforced that location.

Before the group tried to start the engine, we decided we should take it apart and check its condition. It

was the engine from two years ago, hadn’t been started in a while, and possibly needed to be rebuilt.

Upon disassembling the engine, the group determined that all it needed was new gaskets, a new fuel



filter, new fuel lines, and a thorough cleaning. The group assembled the engine, attached the propeller,

filled it up with new fuel, mounted it in a vice, and started it up. After a few attempts, the engine started

and ran without any problems or complications.

Next the group located the servos from the previous years to check if they were in good working

condition. After charging the batteries for both the controller and the servos, the group determined that

all 5 servos were in working order, and that purchasing new ones was not necessary.

After assembling the body of the plane, the group had to decide where and how to mount the servos.

The group placed the aileron servos in each of the wings, with the linkage coming out of a hole

underneath the wing. The servo motor for the flaps was placed in the middle of the top of the fuselage,

as this location was between both flaps and also easily accessible. The other servos were placed in the

body of the fuselage. The transmitter was position underneath the wing itself with the battery to ensure

no change in the center of gravity.

After properly mounting the engine and confirming the functionality of the control surfaces by the

remote, the group took the plane to a small, grass airstrip in Moonachie, NJ. With the aid of an

experienced R/C model pilot, Sean Payne, the plane was tuned before attempting takeoff. The group

then started up the engine successfully, and placed the plane on the runway (grass field). The first few

attempts were unsuccessful, and it was found that the steering and front landing gear needed to be

adjusted. After adjusting those as best as possible at the site, the group attempted to takeoff again.

Steering proved to be a big problem, mainly because the grass was very bumpy and our tires had

problems gaining traction for turning. On our final attempt, the plane reached speed and began to lift.

The plane achieved takeoff, but stability was a problem. At about 6 feet off the ground, the plane rolled

to the right and crashed, breaking the entire tailplane off of the fuselage.



There are a few possibilities for the lack of stability in the plane. Those reasons could include the play

in the joining portion of the wings, an imbalance in the plane, a lack of enough tailplane, the shape of

the tailplane, insufficient control surfaces, construction inexperience, and possibly the engine mounting.

The majority of these shortcomings can be attributed to a general lack of experience in flying model

aircraft; while calculations may have been correct, a lack of first hand knowledge (as opposed to

textbook knowledge) was ultimately responsible for the disfunctionality of the plane.

Conclusion

As the team’s plane did not fly, the group will not be attending the SAE Aero East competition in Fort

Worth, Texas. However, the team will rebuild the tail plane, and continue to make adjustments to the

plane in order to try and get it to fly. It was our overall lack of experience that proved to be our

downfall as the construction was not adequate for flight. Theoretically, from a calculation standpoint,

the plane should fly.

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. The group had some issues getting along,

mainly due to cultural differences and ideological discrepancies between group members. While we had

problems working with each other at points, in the end the ultimate goal of getting the plane to fly united

us.



The group has a few recommendations for future groups. The first would be to consult a local flight

club very early in the design process. The lack of experience was the biggest setback for the group.

Flight clubs have extensive experience and are usually more than willing to offer advice and share their

knowledge. Some of the recommendations they gave were to utilize as large an aspect ratio as possible,

to secure the wing so that very little vibration occurs, to be sure to balance the plane perfectly, to perfect

the engine mount (slight variations dramatically affect performance), to lay the fuel tank horizontally to

avoid head pressure issues, and to use proper axles for landing gear. Next, construct the plane very

early; flight testing is very time consuming, and many adjustments will be necessary before liftoff can

be achieved. Another recommendation would be to order materials very early, as many times there are

delays involved with administrative issues, and backorder issues. Overall, the biggest advice is to

manage time well, and to prepare very early.

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

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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"


26

Appendix D

EES Program I: Drag and Lift Coefficients “Constants" W=25 "Weight" g=32.2 "Gravity" m=W/g "Mass" Rho=.002377 "Density at sea level" C_Lmax=2.018 "Max Lift Coefficient" mew=3.677*10^(-7) "Constant" V=62.33 "Takeoff Velocity" C_Dflapsdown=1.2 "flaps down drag...very conservative" C_Dfrontal=.8 "engine cowl drag...conservative" A_flapsdown=(29.7+14.85)/144 "flaps down frontal area" A_flapsdownLAND=(67.72)/144 "flaps down during landing frontal area" "Dimensions" L_Fuselage=54/12 "Fuselage Length in feet" D_Fuselage=6/12 "Fuselage Diameter in feet" S_WettedFuselageArea=816/144 "Wetted Fuselage Area in feet" T_Wing=.1252 "Wing Thickness" L_WingChord=11/12 "Wing Chord" S_WettedWingArea=880*2/144 "Wetted Wing Area" L_WingThick=1.2 "Wing Airfoil Parameter" R_Wing=1.05 "Constant" L_HTailThick=2.0 "Horizontal Tail Airfoil Parameter" T_HWing=.0902 "Horizontal Tail Thickness"


27

L_HWingChord=5/12 "Horizontal Wing Chord" S_WettedHWingArea=240/144 "Wetted Horizontal Wing Area" R_HWing=1.05 "Constant" L_VTailThick=2.0 "Vertical Tail Airfoil Parameter" T_VWing=.0902 "Vertical Tail Thickness" L_VWingChord=5/12 "Vertical Wing Chord" S_WettedVWingArea=198/144 "Wetted Vertical Wing Area (CHECK)" R_VWing=1.05 "Constant" D_FrontBlockage=5.4/12 "Front Blockage Diameter" R_t=11/12 "Outer Propeller Radius" R_h=0 "Propeller Hub Radius" N=12300 "Propeller RPM" Power=33000*1.7 "Engine Power from Specs" Efficiency=0.85 "Engine efficiency" m_Alpha=-0.008 "Slope- Should it be negative?" ALPHA_ATTACK=.22707 "Angle of attack in Radians" ALPHA_DEFLECTION=.20944 "Angle of Deflection in Radians at Takeoff" ALPHA_DEFLECTIONLANDING=.349 "Angle of Deflection in Radians at Landing" Camber_MAX=.0992 "Maximum Camber" AR=3.2 "Aspect Ratio" A_planform=880/144 "Planform Area" "Drag" Re_Fuselage=Rho*V*L_Fuselage/mew "Fuselage Reynolds Number" C_FFuselage=.074/(Re_Fuselage^.2) "Fuselage skin friction" FR_Fuselage=L_Fuselage/D_Fuselage "?" FF_Fuselage=1+(60/(FR_Fuselage^3))+.0025*FR_Fuselage "Fuselage Form Factor"


28

C_DFuselage=FF_Fuselage*C_FFuselage*S_WettedFuselageArea/L_Fuselage "Fuselage Drag Coefficient" Re_Wing=Rho*V*L_WingChord/mew "Wing Reynolds Number" C_FWing=1.328/(Re_Wing^.5) "Wing Skin Friction Coefficient" FF_Wing=(1+L_WingThick*T_Wing/L_WingChord+100*((T_Wing/L_WingChord)^4))*R_Wing "Wing Form Factor" C_DWing=FF_Wing*C_FWing*S_WettedWingArea/L_WingChord "Wing Drag Coefficient" Re_HWing=Rho*V^2*L_HWingChord/mew "Horizontal Wing Reynolds Number" FF_HWing=(1+T_HWing*L_HTailThick/L_HWingChord+100*((T_HWing/L_HWingChord)^4))*R_HWing "Horizontal Wing Form Factor" C_FHWing=1.358/(Re_HWing^.5) "Horizontal Wing Skin Friction Coefficient" C_DHWing=FF_HWing*C_FHWing*S_WettedHWingArea/L_HWingChord "Horizontal Wing Drag Coefficient" Re_VWing=Rho*V*L_VWingChord/mew "Vertical Wing Reynolds Number" FF_VWing=(1+T_VWing*L_VTailThick/L_VWingChord+100*((T_VWing/L_VWingChord)^4))*R_VWing "Vertical Form Factor" C_FVWing=1.358/(Re_VWing^.5) "Vertical Wing Skin Friction Coefficient" C_DVWing=FF_VWing*C_FVWing*S_WettedVWingArea/L_VWingChord "Vertical Wing Drag Coefficient" C_DLandingGear=3*1.01*2/1440 "Landing Gear Drag Coefficient" A_Frontal=((D_FrontBlockage/2)^2)*pi "Frontal Area for Engine Drag Calculation" C_DEngine=.002 "Engine Drag Coefficient" C_Dinfinity=C_DFuselage+C_DWing+C_DHWing+C_DVWing+C_DVWing+C_DLandingGear+C_DEngine C_Dtotal=C_Dinfinity+C_LFlapped^2/(pi*AR) "Total Drag Coefficient" CDA_TO=C_DFuselage*S_WettedFuselageArea+C_DWing*S_WettedWingArea+C_DHWing*S_WettedHWingArea+C_DVWing*S_WettedVWingArea+C_DLandingGear*(9/144)+C_DEngine*A_Frontal+C_Dfrontal*A_Frontal+C_Dflapsdown*A_Flapsdown


29

CDA_LAND=C_DFuselage*S_WettedFuselageArea+C_DWing*S_WettedWingArea+C_DHWing*S_WettedHWingArea+C_DVWing*S_WettedVWingArea+C_DLandingGear*(9/144)+C_DEngine*A_Frontal+C_Dfrontal*A_Frontal+C_Dflapsdown*A_FlapsdownLAND "LIFT at Takeoff" C_LFlapped=C_Lmax*sin(ALPHA_ATTACK+ALPHA_DEFLECTION+2*Camber_MAX)/(sin(ALPHA_ATTACK+2*Camber_MAX)) "Lift Coefficient with Flaps" "LIFT at Landing" C_LFlappedLanding=C_Lmax*sin(ALPHA_ATTACK+ALPHA_DEFLECTIONLANDING+2*Camber_MAX)/(sin(ALPHA_ATTACK+2*Camber_MAX))


30

Appendix E

EES Program II: The One "Input Values" S_Takeoff= 190 "Takeoff Runway, ft" W=25 "Weight, lbs" C_LmaxTO=3.011 "Flapped Lift Coefficient for takeoff" C_LmaxLAND=3.673 C_L=2.018 "Coefficient at Altitude 1000 ft" AR=3.2 Power=550*1.7 eff=.85 "efficiency" g=32.2 "Gravitational Constant, ft/s^2" e=2.7182818 "Natural Number" rho=.002377 "Air Density at Sea Level" rho_alt=.0023081 "Air Density at 1000 feet for cruise velocity" A_p=880/144 "Wing Planform Area, ft^2" mew_L=.2 "Landing Friction" T_cruise=19.53 alpha_deflectionTO=13*pi/180 alpha_deflectionLANDING=20*pi/180 "Performance" W=m*g "Weight" C_D=.5818 A_pTO=(880-5.07)/144 A_pLAND=(880-11.94)/144


31

CdA_TO= 1.011+((C_LmaxTO^2)/(pi*AR))*A_pTO "C_D times Area for flaps and ailerons down" CdA_LAND=1.7+((C_LmaxLAND^2)/(pi*AR))*A_pLAND "Frank White's method from 'FLUID MECHANICS'" K_TO=.5*rho*CdA_TO V_stall=((2*W)/(C_LmaxTO*rho*A_pTO))^.5 "Stall Velocity, ft/s" V_TO=1.2*V_stall "Takeoff Velocity, ft/s" Y=e^(2*S_Takeoff*K_TO/m) T_1=(K_TO*(V_TO^2)*Y)/(Y-1) "Nicolai's Method from the 'white paper'" D_TO=1.1*(.5*rho*(V_TO^2)*CdA_TO) "Total Drag" a_mean=(g/W)*(T_2-D_TO) "Model" a_mean=(V_TO^2)/(2*S_Takeoff) "Mean Acceleration calculation" T_propeller=Power*eff/V_TO "Landing Calculations" V_Landing=1.7*V_stall V_avgLanding=.7*V_Landing "Landing Velocity" D_Landing=.5*rho*(V_avgLanding^2)*CdA_LAND L_Landing=.5*rho*(V_avgLanding^2)*C_LmaxLAND S_Landing=(V_Landing^2)*(W/g)/(2*(D_Landing+mew_L*(W-L_Landing))) "Cruise Velocity" W=.5*C_L*rho_alt*(V_Cruise^2)*A_p T_cruise=(C_D+(C_L^2)/(pi*AR))*.5*rho_alt*(V_cruise1^2)*A_p

2007 sae aero design east report - stevens institute of · pdf file ·...

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