Aircraft Design Project Sample

8/20/2019 Aircraft Design Project Sample

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2007/2008 AIAA Cessna/Raytheon Design/Build/Fly Competition

Aircraft Design Report

Massachusetts Institute of Technology

Team Concrete


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MIT DBF 2008: Team Concrete

1 EXECUTIVE SUMMARY 3

1.1 Design Overview 3

1.2 System Performance 3

1.3 Design Development 4

2 MANAGEMENT SUMMARY 4

2.1 Organization 5

2.2 Schedule and Planning 5

3 CONCEPTUAL DESIGN 6

3.1 Mission Requirements 6

3.2 Score Analysis 8

3.3 Aircraf t Design Concepts 12

3.4 Configuration Selection 16

3.5

FOM Analysis Results 18

4 PRELIMINARY DESIGN 19

4.1 Design Methodology 19

4.2 Trade Studies and Preliminary Optimization 20 4.3 Payload System 23

4.4 Propulsion System 26

4.5 Aerodynamics 28

4.6 Stability and Control 31

4.7 Estimated Performance 33

5 DETAIL DESIGN 35

5.1 Aircraf t Dimensional Parameters 35

5.2 Aircraf t Structural Character is tics and Capabil it ies 35

5.3

Sub-System Design, Selection, Integration, and Arch itecture 36

5.4 Weight and Balance 40

5.5 Rated Aircraft Cost 41

5.6 Aircraf t and Mission Performance 41

5.7 Drawing Package 41

6 MANUFACTURING PLAN AND PROCESSES 46

6.1 Manufacturing Figures of Merit 46

6.2 Construction Method Selection 46

6.3 Construction Schedule 49

7 TESTING PLAN 49

7.1 Test Schedule 49

7.2 Sub-System Tests and Objectives 49

7.3 Flight Testing 51 8 PERFORMANCE RESULTS 53

8.1 Sub-System Evaluation 53

8.2 Demonstrated Aircraft Performance 57

9 REFERENCES 60

Massachusetts Institute of Technology: Team Concrete Page 2 of 60


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1 Executive Summary

This report describes the design process used by the Massachusetts Institute of Technology (MIT)

Team Concrete to develop an aircraft capable of winning the 2008 AIAA Student Design/Build/Fly

Competition. The goal of the design was to maximize the total competition score, which is a combination

of the report score and three flight mission scores which make up the total flight score.

1.1 Design Overview

The aircraft is essentially a payload compartment with wings. This focus was derived from early

scoring analyses which identified system weight and aircraft loading time as the two key design

parameters. Since the aircraft must be capable of carrying a variety of payload sizes and weights, the

structure required to achieve that objective is potentially the heaviest element of the aircraft. As the

design of the payload system also has a direct impact on loading time, preliminary design focused on the

development of a fast and lightweight payload system capable of meeting restraint requirements with the

minimum aerodynamic features needed to complete lap requirements for the flight missions.

This design takes advantage of a high tensile-strength fabric for the primary payload system

structure. Individual fabric pockets are attached to a central carbon-fiber spar, eliminating the need for a

structural payload bay floor. This innovative fabric payload system is enclosed by a sixty-nine inch span,

twin tractor, low-wing monoplane with tricycle landing gear. The aircraft sits diagonally within the 4 ft x 5

ft planform limits, maximizing aspect ratio and providing additional length for the fuselage fairing, thus

maximizing aerodynamic efficiency.

The aircraft utilizes moldless, foam/fiberglass/carbon-fiber composite construction for the wing, tail

and fuselage internal structure. As the external fuselage takes no structural loads, significant weight

savings were achieved by vacuum-forming a thin, foam shell designed only for aerodynamic loads. Thefoam fuselage fairing has a full-length top hatch which, combined with a low-wing, allows rapid access to

the payload. This payload-focused configuration minimizes the key parameters of system weight and

loading time through its structural efficiency and access to payloads, while providing sufficient

aerodynamic performance and propulsive power density.

1.2 System Performance

The focus on weight in both system design and final manufacturing resulted in an aircraft with a

system weight of 3.02 lbs. A payload loading time of 10-20 seconds is expected when the distance

between the starting area and aircraft ranges from 10-50 ft, respectively, as the rules and FAQ updates

have specified. For the deployment mission, the aircraft lifts off within 20 ft and flies 2 laps in 3 min. 30 s.

when powered by a 4.0 oz propulsion battery pack. During a payload mission scenario, lift off occurs

within 73 ft and the aircraft completes 2 laps in 2 min. 55 s. when powered by a 11 oz propulsion battery

pack. Additional flight vehicle performance parameters which do not directly enter the scoring equation

are provided in Section 5.6.



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1.3 Design Development

During the conceptual design phase, the team focused on analyzing competition rules to select an

aircraft configuration that would maximize competition score. Sensitivity analyses identified system

weight and loading time as key design drivers, with performance in the efficiency-based delivery mission

a secondary factor. A morphological chart was used to enumerate the possible design space of aircraftconfigurations, and a final configuration was selected using a combination of quantitative and qualitative

Figures of Merit (FOMs). The result was the low-wing, fabric payload system concept that was carried

forward into the preliminary design phase.

The preliminary design phase focused on fully developing and refining the details of the design

chosen during the conceptual design phase. The fabric payload system was designed and several mock-

ups were created for full-size testing. The critical aerodynamic design details were determined to be wing

area, aspect ratio, and power requirements at takeoff and cruise. These parameters were optimized

using several in-house MATLAB and Excel-based performance codes, as well as commercial tools such

as XFOIL [1] and AVL [2]. Finally, stability, control, and propulsion system analysis over the entirevelocity range of the aircraft was conducted to further refine the design.

In the detailed design phase, the specific components and manufacturing techniques for the aircraft

were selected, including motors, controllers, batteries, servos, landing gear, and aircraft materials. These

choices were guided by extensive research, as well as the experience and training of the team, which

allowed each component to be built at a low weight and with a high finish quality.

Flight tests verified the predicted performance of the design, providing accurate loading times,

velocities, takeoff distances, and power requirements. There were significant efforts to test the aircraft at

various wind speeds, temperatures, and weather conditions to account for the variation of the expected

environment during a typical late April day in Wichita, KS. The result of this weight and payload-focused

design and testing process was a unique payload-aircraft configuration that maximizes the total

competition score.

2 Management Summary

The 2008 MIT DBF program consists of two teams, Team Concrete and Team Cardinal, which

collaborate to avoid redundant costs and testing. Team Concrete is composed of eight undergraduates,

three of whom are juniors, and five seniors, thus meeting the AIAA Freshman-Sophomore-Junior

competition requirement. The team is led by a Program Manager and then split into three main groups:

Analysis, Design, and Manufacturing. The heads of each group and the program manager form anexecutive board which collaborates to make major design decisions. Due to the small size of the team,

the group members are not necessarily assigned to only one group. The organization of the team is

shown in Figure 2.1.



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Ryan CastoniaPro ram Mana er

Brandon SuarezHead of Anal sis

Dane ChildersHead of Manufacturin

David SanchezHead of Desi n

David SanchezScorin Anal sis

Ryan CastoniaMachinin

Scott ChristopherMachinin

Mikhail GoykmanFabrication

Brandon SuarezFabrication

Riley SchuttFabrication

Martin HolmesFabrication

Dane ChildersPa load

Scott Christopher Aeroshell

Ryan CastoniaCAD

Riley Schutt Aerod namic Performance

Scott ChristopherPro ulsion

Mikhail GoykmanStructures

MITTEAM CARDINAL

Figure 2.1 – Team Organizational Chart

2.1 Organization

The program manager is in charge of the executive board, which is responsible for recruiting new

members, identifying figures of merit, making final design and manufacturing decisions, and ensuring

efficient collaboration with Team Cardinal. Each member of the executive board is in charge of a group

with specific responsibilities.

The analysis group was responsible for creating design trades based on the identified FOMs. They

were also responsible for creating aerodynamic and mission models used to evaluate proposed aircraft

configurations. After deciding on the final architecture, the analysis group became responsible for

component and flight testing. The design group was responsible for providing aircraft configurations to

the analysis group; these configurations were then refined using feedback from the analysis group to

create the final detailed design. After delivering the detailed design to the manufacturing group, the

design group became responsible for the written report. The manufacturing group was responsible for

the production of the aircraft. Their success in building a flying prototype of the proposed configuration by

January 2008 greatly aided with moving forward in the final design process.

2.2 Schedule and Planning

An overall schedule from the beginning of October 2007 to competition was developed by the

executive board. The planned and actual timing of the different phases of the Design/Build/Fly cycle are

shown in Figure 2.2 below.



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Figure 2.2 – Milestone Chart (Red denotes actual timeline)

3 Conceptual Design

This section discusses the details of the conceptual design investigations for the MIT Team Concrete

aircraft. Initial design focused on identifying mission requirements from the competition rules and from a

detailed scoring analysis. Next, a morphological chart of possible aircraft configurations was used to

enumerate the complete design space. Several of the configurations in the design space were eliminated

based on the design team’s qualitative assessments. The remaining configurations were then carried into

a more detailed analysis based on FOM. These FOM were weighted to reflect importance to mission

performance and total flight score. The highest-scoring aircraft configuration, as described in Section 1,

was selected for preliminary design.

3.1 Mission Requirements

Each aircraft must meet a number of payload, structural, performance, and propulsive requirements

for the 2008 DBF competition. The flight competition consists of a single, unloaded delivery flight and two

payload flights. The score of the best performer in each mission normalizes the raw scores of the other

competitors, such that the best performance receives the maximum allowable points for that mission and

other teams receive a corresponding fraction of the possible points. These normalized scores are then

combined using a weighted sum to determine the Total Flight Score.

3.1.1 Payload Requirement

The aircraft must be able to accommodate five possible payload configurations, as seen in Table 3.2,

consisting of various combinations of half-liter, 0.5 lb water bottles and half-size, 1.8 lb bricks totaling 6.8

to 7.2 lbs. Bottles are ballasted with water, include foam collars to limit spacing, and must be carried



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upright within the aircraft. The maximum payload dimensions and possible configurations are given in

Table 3.1. Note the variation in bottle dimensions, as the payload system must securely restrain every

combination of dimensions.

Table 3.1 – Payload Dimensions

Payload Height Max WidthBrick 2.7 in. 4 in. x 4 in.

Bottle 7.6 in – 8.3 in 4 in. x 4 in.

Table 3.2 – Payload Combinations

Number of Bottles Number of Bricks Nominal Weight [lbs]

14 0 7

10 1 6.8

7 2 7.1

3 3 6.90 4 7.2

3.1.2 Flight Requirements

The score for the 2008 competition is determined by Eq 3.1.

TFS RS TS ⋅= (Eq. 3.1)

TS is the Total Score, RS is the report score, and TFS is Total Flight Score. The Total Flight Score term

encompasses the normalized scores from two missions: the Delivery Flight, worth a maximum of 50

points, and the two Payload Flights, worth a possible 50 points each. Each aircraft is also assigned a

Rated Aircraft Cost (RAC ) which is given by Eq 3.2.

BPS W W RAC ⋅= (Eq. 3.2)

Ws is System weight and WBP is Payload Mission Battery Weight. System weight is defined as the weight

of all components of the aircraft minus the propulsion battery weight. RAC is only used in scoring the

Payload Mission. The missions are summarized in Table 3.3 below.

Table 3.3 – Flight Mission Descriptions

Mission Objective Payload Raw Score

Delivery

Fly as many complete laps as

possible in 5 minutes. None Weight Battery Delivery

LapsComplete#

PayloadLoad a given payload combination as

quickly as possible. Fly two laps.

Randomly Assigned

Combination

RAC Time Loading *

1

The aircraft may use a different battery pack for each mission. Each mission score is normalized by

the best team’s score for that mission, with a maximum possible score of 50 points for each flight. This

gives a maximum of 150 points: 50 for the delivery mission and 50 for each completed payload flight.



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3.1.3 Structural Requirements

The maximum aircraft weight may not exceed 55 lbs in any configuration. The aircraft will undergo an

upright wing tip test at maximum payload capacity to simulate wing-root bending moments approximately

equivalent to a 2.5 g load. The aircraft must pass this test without failure of any type.

The payload system must mechanically restrain the bottles and bricks independently of the aircraft

cargo hatch. The restraint system will be tested by inverting the loaded aircraft to present the open cargo

hatch toward the ground.

3.1.4 Geometric Requirements

The assembled, flight-ready aircraft must fit within a 4 ft x 5 ft rectangle in planform view. The aircraft

external surfaces must retain the same external geometry and physical elements for every payload

combination. Payloads may not be exposed to the air stream during flight.

3.1.5 Takeoff Requirements

The maximum takeoff distance for each mission is 75 ft (wheels off the runway). It is important tonote that the field elevation of 1378 ft and ambient temperatures at the competition site will potentially

reduce air density to about 95% of sea level density, depending on temperature and humidity.

3.1.6 Propuls ion System Requirements

The aircraft must use an electric propulsion system. All motors must be commercially available

brushed or brushless electric motors. The battery pack(s) must be commercially available NiCd or NiMH

cells and weigh less than 4 lbs with packaging. The maximum current of all parts of the propulsion

system must be limited by an externally-accessible 40 amp fuse.

3.2 Score Analysis A scoring analysis was performed to identify the most sensitive variables in the total flight score and

assist in the translation of the above mission requirements into design requirements. Ultimately, system

weight and loading time were identified as the most sensitive score variables that would influence the

design. Additionally, analysis revealed the importance of matching battery capacity to the number of laps

flown in the delivery mission and the relative unimportance of absolute flight speed. Table 3.4 provides all

variables used in the analysis.

Table 3.4 – List of Nomenclature

DFS Delivery Flight Score PFS Payload Flight Score

n Number of Completed Laps Lt Loading TimeW BD Delivery Battery Weight W BP Payload Battery Weight

L/D Lift to Drag Ratio W s System Weight

l Lap Length h Cruise height

η Overall Efficiency g Acceleration Due to Gravity

ρ Battery Energy Density



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3.2.1 Delivery Mission Analys is

The Delivery Mission analysis indicated that a competitive design would minimize system weight and

precisely match the delivery battery weight to the energy needed to complete a chosen number of laps.

Using a first order conservation of energy analysis based on an aircraft in steady, level flight at

constant weight, altitude, motor efficiency and power consumption, the Delivery Flight Score (DFS) was

expanded into Eq. 3.4.

( ) ⎟⎟ ⎠

⎞⎜⎜⎝

⎛ −

+==

BD BDs BD W

h

W W gl

D L

W

n DFS

ηρ (Eq. 3.4)

Though approximate, this equation allows preliminary analysis of the relations between variables. It

shows that to maximize DFS, the team needs to minimize system weight and battery weight while flying at

low altitude at maximum L/D. Since the altitude is independent of aircraft configuration and L/D is largely

dependent on the wetted area needed to enclose the payload for a given aircraft weight, the most

important design requirement from the DFS equation is to minimize system weight, with a secondaryrequirement of minimizing drag for a given configuration, which is to be expected.

Using the baseline parameters of a 2500 ft lap length, a 3 lb system weight and 100 points received

on the Payload Missions, TFS versus delivery battery weight and laps completed was plotted in order to

estimate the optimum battery weight. Stored battery energy was assumed to be proportional to battery

weight. Representative values for propulsive system efficiency and battery energy density were

estimated from propulsive systems of previous years at 0.6 and 65mWh/g respectively. Figure 3.1 shows

the effect of increasing delivery battery weight with respect to TFS.

Figure 3.1 – Total Flight Score vs. Battery Weight



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The plot is discontinuous as only complete laps are counted in the scoring equation. Interestingly,

there is only a weak optimum at 2 laps. The peaks in TFS that occur as the battery weight is increased

all generate approximately the same TFS. There is no benefit to carrying extra battery energy to fly an

extra portion of a lap as the extra weight decreases the score, making it important to operate near the

peaks. However, a slight margin should be added because landing short of a completed lap produces the

lowest score. This highlights that the number of laps completed in five minutes is not important; the

important factor is selecting the battery to precisely complete a given number of laps.

3.2.2 Payload Mission Analys is

The Payload Mission analysis indicated that high scores can be achieved by minimizing system

weight and loading time, in that order.

Another first-order conservation of energy analysis based on an aircraft in steady, level flight at

constant altitude, weight, motor efficiency and power consumption, was used to expand the Payload

Flight Score (PFS) equation into Eq. 3.5.

( ) ⎟⎟ ⎠

⎞⎜⎜⎝

⎛ ++

−−

==

D L

lhW W W L

D L

lh

g

W W LPFS

psst

BPst 2

2

1

ρ η

(Eq. 3.5)

This model shows that the team must decrease system weight and decrease loading time in order to

maximize PFS, while still flying at max L/D cruise velocity and a minimum safe height. (W s) 2 appears in

the denominator, making it the most sensitive parameter, meaning that lowering system weight is the

highest priority in the design of this aircraft, with loading time as a slightly lower priority. This can be seen

by the relative change in flight score based on changes in both parameters, shown in Figure 3.2.

Figure 3.2 – Payload Flight Score vs. System Weight and Loading Time



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3.2.3 Normalization and Total Flight Score

As stated before, the total flight score is computed based on the normalized scores from each

mission. In order to assess the effect of score parameters on TFS, a best loading time of 5 sec and best

system weight of 2 lbs were assumed to compute a “best raw score,” which was used to normalize the

rest of the scores. Sensitivity of total flight score to system weight, loading time, and payload battery

weight is plotted in Figure 3.3 with axes scaled to reflect expected parameter ranges. TFS is most

sensitive to system weight, followed by loading time.

Figure 3.3 – Normalized TFS vs. System Weight, Payload Battery Weight, and Loading Time

3.2.4 Design Drivers Conclusion

The scoring analysis resulted in the conclusion that system weight was the most significant figure of

merit, followed closely by loading time. Decreasing system weight tends to also decrease battery weight

required, another parameter in the denominator of both scoring equations. Additionally, matching the

delivery battery to a given number of laps is far more important than the precise number of laps flown.

From this scoring analysis, the two following major design considerations were articulated in order to

focus the conceptual design process:

• System Weight

Decreasing system weight significantly below competitors’ weights is the primary goal. Past

winning payload mass fractions should be used to set aggressive target weights. A winning

design may trade off loading time and some aerodynamic efficiency (i.e. L/D and battery weight)

for decreased system weight. This trade-off between weight and drag should be evaluated.

• Loading Time

The aircraft configuration should facilitate rapid loading of the payload system. Effort should be

made to create a simple, lightweight system with minimal loading steps and components. If

possible, aerodynamic surfaces and internal structures should not impede the loading crew.



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3.3 Aircraft Design Concepts

The first stage in choosing an aircraft concept was the selection of basic payload system

characteristics. This stage was then followed by an enumeration of the design space based on possible

aircraft configurations with varying wing, fuselage, empennage, landing gear, and propulsion

architectures. Table 3.5 summarizes the design space, which when fully enumerated included 768configurations. Using a combination of qualitative reasoning and first-order performance calculations, the

weaker component configurations were eliminated, leaving 8 designs for further analysis.

Table 3.5 – Initial Morphological Chart

Component Types

Wing Monoplane Biplane N-plane Tandem

Fuselage Conventional Blended Lifting

Empennage Conventional V-tail H-Tail Tailless

Landing Gear Tail-dagger Bicycle Tricycle Mono-wheel

Propulsion Tractor Pusher Twin Tractor Twin Pusher

3.3.1 Payload

The payload system is the critical element in the 2008 competition due to its impact on system weight

and loading time. The following design parameters for the payload structure were considered:

• Rigid vs. Conformal

Initial brainstorming resulted in several payload concepts, including racks, removable “quick-loaders”,

and various mechanical locking mechanisms. Ultimately, few of the concepts offered significant

advantages in terms of weight, simplicity or loading time over a fabric pocket design or a rigid box design.

A rigid design could potentially serve as the primary aircraft structure, though the requirement of an

additional payload restraining hatch in addition to an external fuselage hatch was considered an

unfavorable weight penalty. A “soft” fabric restraint system, closed with a draw-string, was ultimately

chosen for its low weight and ability to conform to the wide variety of payload dimensions.

• Loading Direction

Three options for the loading direction were considered: side loading, top loading, and bottom

loading. The side loading and bottom loading configurations potentially provide a weight advantage by

circumventing a complete overturning of the aircraft during the flip test, which would require additional

restraints. However, these systems required significantly higher loading times. The top loading

configuration best capitalizes on the normal top-down motion required to load a small RC aircraft and was

kept for further analysis.

• Payload Configuration

The five payload configurations could be arranged in many ways to maintain a near constant center

of gravity. Aircraft drag considerations (frontal area vs. wetted area) and the 4 ft x 5 ft planform

requirement (space limitation when trying to fair in the payload system) resulted in the three possible

configurations shown in Figure 3.4. The 2 x 7 configuration was ultimately selected due to its ability to be



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primarily supported by one central spar running lengthwise between the pockets, unlike the other two

configurations which require multiple spars and thus increase system weight and complexity.

3-8-3 2 X 7 2 X 6+2

Figure 3.4 – Possible Payload Configurations

In order to minimize system weight and loading time, the selected payload system was a fabric, top-

loading, 2 x 7 draw-string closed configuration as shown in Figure 3.5.

Figure 3.5 – Fabric Payload System Concept

3.3.2 Wing

Typically, the simplicity and performance per weight of the monoplane would make it the frontrunner.

Despite this, the span and aspect ratio limitation from the 4 ft x 5 ft planform made a multi-wing aircraft anattractive option. However, the tandem wing was eliminated because it provided few if any benefits

compared to the other multi-wing configurations while potentially adding weight (due to a larger section of

structural fuselage) and risk (due to stability and lift distribution issues). The N-plane, with N>2 wings,

was eliminated because of downwash and venturi interference, reduced wing efficiency, and doubts

about the team’s ability to construct sufficiently light wings to realize the benefits of lower wing loading.

The monoplane and biplane were retained for more detailed analysis, with the understanding that the bi-

plane would require the top wing to be hinged or split to facilitate the top loading payload system.

Table 3.6 – Wing types

Monoplane Biplane N-plane Tandem



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3.3.3 Fuselage

While the lifting fuselage could potentially reduce wing loading, it was eliminated because of the

difficulty of executing low-weight construction and excessive airfoil thickness due to payload height and

planform constraints. Conventional and blended fuselages were retained for more detailed analysis.

Table 3.7 – Fuselage Types

Conventional Blended Lifting

3.3.4 Empennage

The H-tail was initially considered to increase the effectiveness of the horizontal control surface

through endplate/winglet effects due to tail length limitations. It was eliminated due to the weight ofmultiple vertical tail surfaces with extra control servos. The V-tail was not considered; the area required

to achieve control equivalent to a conventional tail resulted in no savings in system weight. The

conventional and tailless configurations were retained for more detailed analysis; the former for its low

risk and the latter for the possible weight advantage if combined with a reflexed wing airfoil.

Table 3.8 – Empennage Types

Conventional V-tail H-tail Tailless

3.3.5 Landing Gear

While ground handling is not explicitly emphasized in this year’s competition, the threat of strong

crosswind gusts and the configuration of the payloads eliminated the single wheel and bicycle landing

gear options. Based on pilot input regarding the limited take-off length and ground stability, a steerable

tricycle landing gear type was retained for more detailed analysis.

3.3.6 Propulsion

A sample of commonly available electric motors showed a clear trend – the smaller motors

consistently had higher power density, as much as 250% difference over their larger cousins. Given the

importance of system weight in total flight score, this finding was used as the basis of eliminating both

tractor and pusher single motor configurations. Additionally, the twin pusher configuration was discarded

due to structural (wing thickness at trailing edge) and motor cooling considerations. Thus the twin tractor

configuration was retained for more detailed analysis.



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Table 3.9 – Propuls ion Types

Tractor Pusher Twin Tractor Twin Pusher

Given the choice of twin motors, a decision had to be made regarding the use of a single or dual pack

(in parallel) battery configuration. A survey of available battery cells showed significant energy density

peaks around 1500 mAh and 2000 mAh, suggesting the use of a single pack would result in a lighter

propulsion system. However, a dual configuration would potentially require less current draw from each

pack, increasing effective capacity. Ultimately, the single pack configuration was selected to minimize

weight.

3.3.7 Final Morphological Chart

Table 3.10 is the revised morphological chart in which the component types eliminated in the previous

section were removed from consideration. The table features 2 wing types, 2 fuselage types, 2

empennage types, and 1 propulsion type; thus there were 8 possible aircraft configurations to investigate

in more detail.

Table 3.10 – Revised Morpho logical Chart

Wing Fuselage Empennage Propulsion

At this point each configuration was qualitatively assessed with particular emphasis on:

• System weight

• Loading time• Manufacturability

• Design risk (i.e. lack of previous flight experience)

• Stability & controllability

These criteria were used to narrow the design space to the four configurations shown in Table 3.11.



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Table 3.11 – The four configurations analyzed us ing f igures of merit

MonoplaneConventional FuseConventional Tail

Twin Tractor

MonoplaneConventional Fuse

TaillessTwin Tractor

BiplaneConventional FuseConventional Tail

Twin Tractor

Blended Wing BodyTwin Tractor

3.4 Configuration Selection

The selected configurations were analyzed with four qualitative and quantitative FOMs. The

qualitative FOMs – Stability and Control and Manufacturability – were assigned a score between -1 and 1.

The quantitative FOMs – System Weight and Loading Time –made use of performance estimations.

Each FOM was weighted based on its importance to strong performance at the competition. The sum of

the weight factors was 100.

3.4.1 System Weight

The most important quantitative FOM was System Weight, due to its strong score influence. A

spreadsheet was developed to estimate wing area and power requirements for each configuration. Using

weight fractions from past MIT aircraft, system weights were estimated as shown in Table 3.12.

Table 3.12 – Estimated System Weights

Monoplanew/ Tail

MonoplaneTailless

Biplanew/ Tail

Blended Wing Body

System Weight [lbs] 2.8 2.7 3.2 3.0

3.4.2 Loading Time

Loading Time is mainly influenced by aircraft configuration, as competition history has shown that

time required to move payload elements from starting locations to the aircraft is roughly constant among

teams. A loading time figure of merit was assigned to each configuration, based on the following:

• Wing/hatch interaction

• Aircraft component interference

• Use of natural loading movements

Table 3.13 shows the results of these estimations.

Table 3.13 – Estimated Loading Times

Monoplanew/ Tail

MonoplaneTailless

Biplanew/ Tail

Blended WingBody

Loading Times [sec] 10 10 20 15

3.4.3 Stability and Control

In this year’s competition, stability and control are crucial as competitive aircraft must fly multiple

missions with varying weight distributions. The possible payload configurations have a maximum



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horizontal CG variation of 0.5” and roughly triple the weight of the unloaded aircraft. A stability and

control FOM was qualitatively assigned to each configuration, based on the following factors:

• Robust longitudinal stability with CG variation

• Lateral and directional stability

• Ground handlingThis FOM was assigned a weight factor of 10 because of the role of flight characteristics and ground

handling in preventing crashes. Table 3.14 shows each configuration’s assigned score:

Table 3.14 – Stabili ty and Contro l FOM Criteria

Assigned Score Configurat ion Character istic

-1 Exhibits weak performance with respect to criteria

0 Exhibits moderate performance with respect to criteria

1 Exhibits strong performance with respect to criteria

3.4.4 Manufacturability

Manufacturability is defined as the feasibility and complexity of fabricating a concept. While the

quality of the aircraft design plays a large role in determining final performance, the execution of the

design also plays a significant role. As such, the team was concerned with choosing a competitive design

that was feasible to execute with a low system weight and without excessive time. A manufacturability

FOM was qualitatively assigned to each configuration, considering the following factors:

• Structural complexity

• The team’s prior experience in building techniques

• Required time and money

Table 3.15 shows how the FOM scores were assigned to the configurations. This FOM was assigned a

weight factor of 20 because of its influence on system weight and limited project time.

Table 3.15 – Manufacturabil ity FOM Criteria

Assigned Score Configurat ion Character istic-1 Little or no prior experience in required fabrication techniques AND

Structurally complex design

0 Prior experience in required fabrication techniques ORStructurally simple design

1 Prior experience in required fabrication techniques ANDStructurally simple design

3.4.5 Mission Performance

To estimate mission performance, simple foam mockups of the payloads and fuselage configurations

were developed to test high-level performance of the payload loading configuration. Delivery and

Payload mission flight scores of each configuration were estimated using a mission profile simulation,

which is discussed in Section 4.1, Preliminary Design.



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3.5 FOM Analysis Resul ts

The results of the FOM analysis are shown in Table 3.16. The Total FOM represents the weighted

sum of the normalized performance of each configuration and the weighting factor of each FOM.

Table 3.16 – FOM Results

S y s t e m W e i g h t

D e p l o y m e n t F l i g h t

L o a d i n g T i m e

A v g P a y l o a d

F l i g h t S c o r e

T o t a l F l i g h t S c o r e

N o r m a l i z e d S c o r e

S t a b i l i t y & C o n t r o l

F O M

M a n u f a c t u r a b i l i t y

F O M

T o t a l F O M

ConfigurationWeighting Factor 70 10 20 100

Monoplane w/ Tail 2.8 45 10 45 135 1.42 1 1 129

Monoplane Tailless 2.7 40 10 45 130 1.37 -1 1 106

Biplane w/ Tail 3.2 35 20 30 95 1 1 0 80

Blended Wing Body 3.0 40 15 35 110 1.16 0 -1 71

3.5.1 Initial Configuration Selection

As Table 3.16 shows, the two monoplane configurations were clearly the strongest performers due to

their low system weight and loading time. The Total FOM results of the two monoplane configurations

were quite close, which is somewhat expected since they build on similar concepts and essentially only

differ on the tail component. Thus the team decided to take a closer look at the two monoplane

configurations, with special consideration to the high-risk areas of each design.

3.5.2 Final Configuration

The tailless monoplane design, despite stability considerations, did have some advantages due to its

low system weight and thus high predicted payload mission scores. Closer analyses of the tailless

monoplane revealed that, in order to achieve the wing area required for takeoff and still remain within the

4 ft x 5 ft planform, a significantly larger chord and thus lower aspect ratio would be required as compared

to the monoplane with tail. This is due to the reduced efficiency of the reflexed airfoil required for a

tailless configuration. The subsequent increase in wing weight and drag negatively affected L/D,

diminishing the tailless monoplane’s competitiveness; hence, the tailed monoplane design was chosen.

The concept sketch for the selected design is shown in Figure 3.6. The design uses two small, high

power/weight brushless motors running on a single battery pack for propulsion and a steerable tricycle

landing gear for ground handling. Instead of a conventional fuselage, the aircraft utilizes an internal

frame to support a lightweight fabric payload system and non-structural aerodynamic fairing with a top-

loading hatch, thus minimizing system weight and loading time.



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Figure 3.6 – Final Concept Sketch

4 Preliminary Design

4.1 Design Methodology

In order to quantify system weight and loading time requirements and to estimate competitors’

performance, the team researched the payload/system weight fractions, wing loading, and loading hatch

orientations of all first place DBF designs from the past three years. The resulting weight targets, wing

sizing estimates, and payload system requirements are discussed in the following sections.

Additionally, the team searched for current technological opportunities that would provide a

competitive weight advantage. From this research, the team identified the highest power to weight ratio

electric motors and highest energy density batteries available. Tests of numerous motor/prop/battery

systems allowed the team to optimize aerodynamic surfaces to match the physical capabilities and

efficiencies of specific lightweight, high-performance motors. The team also researched the lightest RC

construction methods available in order to design an efficient, manufacturable structure. However, the

most significant result of the team’s research was the discovery of a high strength to weight fabric that

would enable the construction of a fabric payload carrier. The details of this design feature are provided

in the payload design section.

Once the team established basic flight performance and identified viable high-performance propulsion

systems and construction methods, it began the process of sizing aerodynamic surfaces and structural

components to meet mission requirements. These primary mission requirements were the 75 ft. take-off

distance, 4 ft x 5 ft planform compatibility, and sufficient range and controllability to fly two laps with or

without payload. The iterative preliminary design process used a combination of custom-developed

multidisciplinary optimization codes, commercial software, and hands-on testing to predict mission

performance, optimize aerodynamic loading, estimate aircraft stability, and size structural components.

The design flow is shown in Figure 4.1.



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Design Research

High PerformancePropulsion Systems

Power output estimates

Competitive ConfigurationsWeight fractions, payloads,airfoils, and wing loadings

Manufacturing Techniques Advanced materials, historical

weight estimation

Payload System DesignStructure and Weight

Estimate

Preliminary A ircraft Sizing

Aero Design

Mission Requirements and Score Analysis

Model Mission PerformanceIdentify design deficiencies,

Iterate Design

Figure 4.1 – Flowchart of Preliminary Design and System Optimization

4.2 Trade Studies and Preliminary Optimization

Preliminary design of the aerodynamic surfaces was performed assuming a 2 x 7 payload

arrangement. The conceptual design scoring analysis indicated that maximizing aircraft efficiency, or

L/D, would play a significant role in battery weight. It was decided to perform an analysis of total energy

consumption over each mission in order to compare preliminary designs and explore the trade-off

between drag and weight. The team began by creating a series of models to estimate the weight of wing

and tail surfaces, size tail surfaces based on wing span and tail arm lengths, and relate planform

limitations to possible aircraft dimensions. Initial wing sizing was performed using previously successful

DBF wing loadings as initial conditions. The nearby design space was explored numerically using the

MATLAB model described in the following sections.

This preliminary aircraft optimization resulted in a basic aircraft geometry which served as a starting

point for the design and refinement of the payload structure, propulsion system, detailed aerodynamics,

and stability characteristics.

4.2.1 Histor ical Research

A survey of past winning teams revealed that the Oklahoma State University Black Team of 2006 had

the highest relevant payload to system weight ratio of any team in the last three years. The 2007 MIT

team’s weight fraction, though lower, was skewed by the lack of a fuselage in their total system weight.

OSU’s system weight of 3.79 lbs for an 8 lb payload gives a system/payload mass fraction of 0.47, which

applied to this year’s payload of 7.2 lbs, gives a competitive system weight of 3.38 lbs. The team used



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this system weight in preliminary analyses and set an internal goal of 10% improvement, giving 3.05 lbs.

as a target system weight.

Successful teams have traditionally undersized wing area based on site wind assumptions. The

recorded wind speed and temperature at the site location in Wichita, KA for each of the last 3 years

during the week of the competition was found at Wundergound.com and averaged over daylight hours in

order to identify an appropriate estimate for headwinds and air density [3]. Based on wind data, runway

altitude, and temperature, conservative estimates were found to be 10 mph and 95% of sea-level air

density, respectively. However, further research of past team reports and websites revealed that several

teams in the last five years had wasted take-off attempts during brief, unrecorded periods of calm wind.

Based on this knowledge, but recognizing that downsizing the wing may provide a competitive weight

advantage, the team made the decision to size the wing for a zero-wind 75 ft. take-off with no margin.

Additionally, a wing-loading of 2.5 lbs/ft2 was identified as a viable starting point for preliminary

optimization based on historical wing areas and predicted wind speeds.

4.2.2 Primary Mission Model

Flight missions were modeled using a 2 degree-of-freedom mission simulation written in MATLAB.

The simulator featured a flight derivatives engine which calculated position, velocity, energy consumption,

and lap times. The purpose of this model was to predict mission performance and estimate battery size

through integration of forces on a simulated aircraft. The inputs to this engine are three model files, which

captured the relevant parameters of the mission, aircraft, and competition site. These files include:

Mission profile model file: Contains information on the sequence of activities (e.g. takeoff

distance, turn radius, level flight distance) in a given mission.

Aircraf t configurat ion f ile: Contains information on vehicle weight, lifting surface dimensions,

aerodynamic coefficients (lift and drag), and propulsion system information.

Site conditi ons file: Contains information about air density and wind conditions.

The course was modeled using four distinct mission segment types – takeoff, climb, turning, and

cruise. No ground operations were modeled. The payload loading time is estimated from the aircraft

configuration and input separately. Additionally, landing ground roll was not modeled since both missions

are essentially completed in the air, with only a successful landing required to confirm score. All missions

are modeled using the throttle settings and lift coefficients shown in Table 4.1.

Table 4.1 – General Flight Mission Segment Profi le

Mission Segment # inmission

Length[ft]

CL Throttle Notes

Take-off (at rotation) 1 Dto CL, max MaxDto = takeoff distance

(calculated)

Climb 1 500-Dto CL, max climb Max

Level Cruise 7 500 CL, max L/D T=D

Turn (180° each) 6 * CL, max L/D T=D*Length given by user-specified turn radius



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The details of each segment model were as follows:

Takeoff : A rolling coefficient of friction of 0.03 was assumed, based on empirical data for plastic

on concrete. The take-off roll continued until the aircraft reached 110% of the stall speed, with a

2 second rotation added on.

Climb: Given the takeoff distance calculated in the takeoff segment, the residual distance to the

first turn was calculated. A constant climb rate was calculated based on attaining 75 feet of

altitude before the first turn.

Cruise: In this segment, altitude was assumed constant with throttle set to equate thrust and

drag at cruise velocity and CL.

Turns: The turns were also modeled with constant radii, velocities, and altitudes. Thus a

constant load factor was assumed, with CL limited to CL, max as specified by the aircraft

configuration file. The 360° turns were modeled as two back-to back 180° turns.

The uncertainties of this model are primarily related to accurate drag prediction and the actual

operation of the aircraft by a human operator. In actual flight, turns are often made at a less than ideal

radius and climb-out may be made at a non optimal point on the aircraft power curve. Additionally,

varying wind conditions or aircraft instability may result in unplanned side-slips, turns, or climbs, all of

which increase power consumption and are not modeled. Finally, the drag model of the aircraft in the

configuration file must be accurate for the total energy consumption to be correct. The preliminary drag

model is based on skin friction estimates with form-factors and has proven sufficient for initial sizing and

head-to-head comparisons of different designs. However, the team treated absolute energy consumption

estimates as lower-bounds and scheduled flight testing to validate model performance.

4.2.3 Aspect Ratio and Wing Loading Optimization

Initial wing sizing was conducted using a MATLAB script which takes in several geometric, propulsive

and aerodynamic constraints, derives wing and tail surface geometry, and then calculates the drag of the

aircraft at cruise, climb and take-off conditions. This aircraft configuration model is then fed into the

previously described mission model.

Code parameters were varied to explore the design space near a 3 lb system weight, 5 ft span

monoplane. The code accounts for structural weight based on a constant curvature wing-bending model

assuming carbon-fiber composite construction. Span was varied within the 4 ft x 5 ft planform constraints

to assess parasitic drag, induced drag, and weight trade-offs while taking into account Reynolds number

effects. Drag is estimated based on total aircraft surface area, using flat plate skin friction coefficientsand form-factors based on the thickness of the aerodynamic surfaces and basic fuselage and landing

gear geometry [4]. Additionally, the model contains drag polars from several low-Reynolds number

airfoils described in Section 4.5 [5]. Tail geometry is calculated from historical tail volume coefficients,

which are described in Section 4.5, and the geometric constraints of the box. The propulsion system was

modeled with a series of thrust-velocity and efficiency-velocity curves from preliminary motor testing



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described in Section 4.4. The model was limited to 6 lbs maximum static thrust based on the weight

penalty for using motors capable of producing greater thrust.

Each aircraft geometry was run through a simulated mission, and the wing area, span, sweep, and

planform orientation were iterated in 1 in2, 1 in, 0.5 degree, and 0.5 degree increments, respectively, until

total flight score was maximized. Taper ratios and tail sweep were set to be dependent on planform

constraints. The preliminary optimized aircraft takes advantage of the planform diagonal to increase wing

span and total aircraft length, as shown in Figure 4.2. From this initial optimization, further refinement of

the wing, airfoils, structure, and propulsion system was performed with component-specific design tools.

Figure 4.2 – Preliminary Aircraft Parameters and Planform View (dimensions in inches)

4.3 Payload System

After initial wing sizing, a viable payload system was designed to allow sizing and engineering of the

fuselage for later use in more detailed aerodynamic design. Physical testing of payload system

prototypes was a central aspect of payload preliminary design.

4.3.1 Support and Restraint Design

Initial bench-top experiments, as shown in Figure 4.3, and a review of past winning payload system

designs led to internal requirements for a top-loading system with single-step actuation, in order to

minimize loading time. It was determined that loading times could be made comparable to rigid box

designs as long as the soft designs had a minimal structural framework to keep the system open while

loading. In addition, the minimum opening through which a 4 in x 4 in payload item could be loaded

quickly and efficiently was determined to be 4.5 in x 4.5 in.

Wing Span 5 ft. 9 in

Wing Area 4.28 ft2

Aspect Ratio 7.72

Root Chord .993 ft

Taper Ratio .51

Horizontal Tail Span 24 in

Horizontal Tail Area 0.91 ft2

Required Max CL 1.44

Cruise CL (max L/D) .85



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Figure 4.3 – Payload System Prototype

During initial materials research, Cuben Fiber high-strength sailcloth with an area density of 0.33

oz/yd2 was identified as an ideal material for our fabric payload design. After initial mock-ups, a

preliminary design of a Cuben Fiber soft payload system was created with 14 individual pockets and a

carbon-fiber tube framework to keep the system open during loading. The individual pockets ensure that

payload items are visibly separated and held upright at all times, conforming to the competition rules.

The design uses a simple drawstring closing mechanism to restrain the payloads while inverted and all

seams are secured using a polyurethane adhesive and Nylon stitching. Figure 4.4 shows the initial

Cuben Fiber payload system used for preliminary payload loading and flip tests. Total weight of the

Cuben Fiber, central spar, and carbon rods was estimated at 4.7 oz.

Figure 4.4 – Cuben Fiber Payload System

All possible variations in CG due to variation in payload weights were calculated for use in aircraftstability calculations. Figure 4.5 shows the optimized layouts. These configurations have a maximum CG

shift of .42 in from front to back and .09 in from side to side. The act of constraining the bricks to the four

specific brick pockets aids in minimizing CG shifting while decreasing loading complexity between the

different payload configurations.



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Figure 4.5 – Payload Configurations

4.3.2 Fuselage Design

Once the basic wing, tail surface, and payload system geometry were defined, a complete

Computational Fluid Dynamics model of the aircraft was created to inform the preliminary design of the

fuselage. Using STAR-CCM+ software, an integrated software package with an extensive selection of

turbulence models, the fuselage was redesigned and analyzed iteratively until no separation was

predicted. Figure 4.6 provides an example of the simulated flow. Details on initial CFD drag predictions

and wind-tunnel testing of physical models are provided in Section 4.7 and Section 8.1, respectively.

Figure 4.6 – Fuselage Flow Visualization

Fuselage manufacturability tests were performed concurrently with the preliminary fuselage

computational design. Manufacturing research early in the design process highlighted the possibility of

using vacuum-formed Depron foam as a fairing material. The team built a male plug of the preliminary

fuselage design and performed numerous fabrication trials with varying thicknesses of Depron and other

varieties of polystyrene foam. Figure 4.7 shows initial shaping of a foam nose section over a male mold.



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Figure 4.7 – Fuselage Nose-Cone Manufacturing Trail

4.4 Propulsion System

Preliminary motor testing provided baseline thrust and power values for use in the initial wing sizing

and optimization. Further testing was carried out to match propellers, establish maximum thrust

performance, and fully document motor efficiency for use in later mission simulations.

4.4.1 Motors and Electronic Speed Control lers

A survey of electric motor brands and models was conducted to compile a list of the highest power to

weight motors available; the compiled list is shown in Table 4.3.

Table 4.3 – Power to Weight Comparison of Commercially Available Motors

Motor Power Output [W] Weight [g] Power/Weight [W/g]

LittleScreamers Park Jet 185 25 7.4

LittleScreamers Purple Peril 165 25 6.6

LittleScreamers DeNovo 128 25 5.12

JustGoFly 450FT 245 60 4.08

JustGoFly 300DF 115 30 3.8

JustGoFly 400ST 110 38 2.9

Extreme Flight Torque 34/1520 81.4 29 2.8

AXI 2212 150 57 2.6

JustGoFly 500SH 400 62 6.45

JustGoFly 500T 250 62 4.03

JustGoFly 500XTF 250 62 4.03

JustGoFly 500XT 250 62 4.03

All of the highest power to weight ratios were from the smallest (~25 gram) class motors and initial

testing demonstrated that these motors produced a maximum of 1.7 lbs static thrust. However, analysis

of these motors in the mission model showed that take-off field length requirements imposed enough

additional wing area to overwhelm the weight savings.

The JustGoFly 500 range of motors was selected for further testing of maximum thrust capabilities.

Ultimately, the 500XT was chosen as the best candidate for extensive testing with multiple propellers over

the aircraft velocity range. Initial tests showed that both the 500XT and the 500XTF were able to produce

3 lbs of static thrust; however the XTF motor showed a greater loss in thrust over the lifetime of the motor.



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Additionally, research showed that the highest power to weight ratio JustGoFly motor, the 500SH, was

designed for high RPM applications in R/C helicopters. Thrust output was below the 3 lbs per motor

produced by the other motors in the 500 series and a suitable gearbox for the 500SH was not available

for testing. Finally, the 500T motor was only able to produce 2.5 lbs of thrust while drawing 36 amps at

12 volts. Two parallel motors pulling 36 amps each would be beyond the safe amperage range for the 40

amp safety fuse, so testing of the 500T was not carried out for higher voltages.

4.4.2 Propeller

The main FOMs for propeller selection were:

• Takeoff Thrust: Due to the heavy payloads and moderate wing loading, a large amount of thrust

is needed to meet the 75 ft takeoff distance requirement.

• Cruise efficiency: Low efficiency due to improper matching of the propeller to motor speed

increases battery size, which increases RAC and limits Delivery Mission performance.

The weighting of each FOM is dependent on which mission is being considered due to the difference in

total aircraft weight between the Delivery and Payload mission. Propellers were selected independently

for each mission.

Delivery Mission Propeller Selection

The Delivery Mission requires that the aircraft fly with the smallest battery pack possible and with the

number of laps closely matched to the capacity of the chosen battery pack. To minimize the battery

weight, a pack with the fewest cells would be optimal. The mission takeoff model calculated 1.0 lb of

static thrust was needed for takeoff at 24 ft/s, with 0.75 lbs needed for cruise at 30 ft/s. Propellers were

then tested for maximum thrust at these two speeds at a range of voltages. The minimum voltage that

provided the required thrust was 6.0 volts. The optimum propeller size was found to be in the 10 to 12 in

range with a very high pitch. This is due to the low RPM of the motor requiring a higher pitch from the

prop to provide necessary thrust.

Of the 6.0 volt runs, the APC 11x10 and 12x12 passed the minimum average thrust requirement.

Both propellers drew approximately 7 amps while providing the same thrust and were retained for

extensive testing with selected battery packs during detailed design.

Payload Mission Propeller Selection

The Payload Mission requires maximizing low speed thrust to meet the 75 ft takeoff requirement. The

takeoff model determined that 6 lbs of static thrust was needed at takeoff due to the higher takeoff weight.

Large, low pitch props would be needed to provide the necessary low speed thrust.

The full range of APC propellers were considered for the delivery mission, though focus was

concentrated on the 10 in to 12 in propellers. It was found that the recommended 11x5.5 propeller for the

JustGoFly 500XT motor operating at 14.4 volts performed best, providing 3.1 lbs of static thrust. The

slightly higher pitch 12x6 propeller performed similarly, but drew 21 amps at approximately 270 watts.

These values were experimentally found to be below, but near, the burn out limit of the 500XT motor.



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Both propellers were retained for further flight testing. Section 8.1 discusses propulsion system testing

and verification using simulated wind-tunnel mission flight profiles.

4.4.3 Batteries

The primary mission model predicted an average electric power consumption of 360 watts at a cruise

speed of 48 fps, resulting in lap times just over one minute. The delivery mission would be completed at

a slower cruise speed of 28 fps and average power draw of 70 watts, requiring approximately 85 sec. per

lap. Batteries of candidate sizes were compared by their energy densities, as shown in Figure 4.7.

Power Density Comparison

59.6 62.9

73.6 77.1

69.9

60.0

47.9

73.4

65.4

0.0

10.0

20.0

30.0

40.0

50.0

60.0

70.0

80.0

90.0

G P 1

1 0 0

C B P

1 1 5 0

I B 1 4 0 0

E L I T E 1

5 0 0

C B P 1

6 5 0 A

A

H R 1 7 0

0 A U P

H R 1

9 5 0 F A U P

E L I T E 2

0 0 0

G P 2

0 0 0

m W h p e r g r a

Figure 4.7 – Battery Power Density Comparison

From the energy consumption estimated in Section 4.1 and simulations to be discussed in Section

5.3, it was determined that a minimum of five 1.2 V cells at 900 mAh capacity and twelve 1.2 volt cells at

1300 mAh capacity would be needed for the delivery and payload missions, respectively. For the delivery

mission a GP1100 NiMH pack was determined to have the lowest weight, due to the lower weight per

1100 mAh cell in comparison to the larger capacity cells. The Elite 1500 NiMH cell stands out as having

the highest energy density in the 1500mAh range. This cell is also able to handle a 40 amp current drain,

so it was selected for testing for use in the Payload Mission.

4.5 Aerodynamics

Building on the baseline geometry established in section 4.1, aerodynamic optimization focused on

the selection and design of airfoils to improve cruise and take-off performance and evaluation of the entire

wing-tail system using vortex-lattice methods to optimize lift-distribution and predict stability.

4.5.1 Airfoil Optimization

Energy Density Comparison

The optimum airfoil for the aircraft is able to provide high lift (CL ~ 1.5) during takeoff and landing

portions of the flight while still having low drag during cruise (CL~.85). A number of airfoils were

considered and were divided into three groups: low, medium, and high lift. High drag penalties at cruise

conditions caused the team to rule out several high-lift airfoils (Max CL ~ 1.7), as the advantage of high



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CL’s would be offset by higher cruise drag and the need for added propulsion battery weight. As

determined by the initial wing optimization, the SD7043 airfoil, with a maximum CL of 1.5, provided the

highest performance of the Selig-Donovan, NACA, MIT-designed, and Douglas LA-series in the airfoil

database. Using the SD7043 performance as a starting point, the team redesigned two series of new

airfoils with maximum L/D at CL~.8 and a max C

L of 1.5 based on interpolation of the LA203 and an

internally designed, BA10 airfoil. The XFOIL drag polar of the resulting airfoils, the BAFT2 and SOFT500,

are shown below in Figure 4.11, along with the SD7043 used in preliminary optimization.

Figure 4.8 – Airfoil Drag Polar Comparison

The drag polar in Figure 4.8 shows that the drag bucket of the BAFT2 provides comparable cruise

drag, but significantly better high lift performance when compared to the SD7043. The SOFT 500 max

L/D occurs at a CL of .85, and provides the lowest cruise drag overall by a 4% margin. However, mission

profiles indicated that the BAFT 2 provides the highest performance, providing a 9% improvement in

energy consumption with no wing weight penalty when compared to the original optimum SD7043 airfoil.

Figure 4.9 – BAFT 2



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4.5.2 Wing and Tail Optimization

In order to estimate the performance of the 3D configuration, a vortex lattice model of the wing and

tail surfaces was created within Mark Drela’s Athena Vortex Lattice (AVL) code, as shown in Figure 4.10.

This model, combined with viscous calculations from XFOIL, allowed estimation of pitching moment

changes with angle of attack in order to estimate neutral point locations and ensure stability during take-

off, cruise, and stall conditions.

Figure 4.10 – Wing-Tail AVL model

Additionally, Trefftz Plane analysis was used to update lift distribution and span efficiency factors

estimated in the MATLAB mission model, and wing washout was adjusted to improve sectional CL

distributions and to reduce the risk of tip stalls. Figure 4.11 shows the lift distribution at cruise conditions.

Five deg of decalage was added to limit elevator deflection at cruise conditions and reduce trim drag.

Figure 4.11 – Trefftz Plot (Lift Distribution at cruise, CL = .9)



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Winglets were considered to increase the span efficiency and decrease cruise drag. A separate AVL

analysis determined that a 0.03 - 0.05 increase in span efficiency was possible. However, the two inch

reduction in span, increased weight, and skin friction from the additional surfaces outweighed the benefit.

4.5.3 High Lift Devices

To meet takeoff CL requirements of 1.5, the wing requires a full-span flap. Preliminary flap sizing was

performed using XFOIL, then final flap areas were determined using AVL to account for 3-D effects. Flap

deflection was limited to 10 deg as larger deflections resulted in excessive separation. Final flap areas

were determined to be 30% of the total wing area.

4.5.4 Empennage

Initial tail surfaces were sized within the mission model to meet static stability requirements based on

tail volume coefficients outlined below [6]. In later refinement of the preliminary design, tail surface areas

were adjusted using AVL to satisfy minimum trim drag conditions.

Horizontal Stabilizer

The horizontal tail volume coefficient (V h) is a measure of horizontal stabilizer effectiveness, and is

defined by Equation 4.1. For sufficient pitch authority, V h > 0.30 is required.

Sc

lS V hhh ≡ (Eq. 4.1)

Wing area and chord (S and c ) are given by the wing geometry, and the tail moment arm (l h) is restricted

by the planform geometry. Using these parameters, Equation 4.1 can be solved for Sh to give an

approximate horizontal stabilizer area of 0.9 ft2.

Vertical Stabilizer

Like V h, the vertical tail volume coefficient (V v ) characterizes the effectiveness of the vertical

stabilizer, and is defined by Equation 4.2. For sufficient yaw damping, a V v > 0.02 is necessary.

Sb

lS V vvv ≡ (Eq. 4.2)

Again, l v , S, and b are specified by the wing and tail boom geometry, and using these parameters

Equation 4.2 can be solved for Sv to give a vertical stabilizer area. Test cases were run in AVL at both

cruise and takeoff conditions. In both cases, the aircraft exhibited positive spiral stability, requiring no

adjustment to the vertical stabilizer area of 0.58 ft2 given by Equation 4.2.

4.6 Stabili ty and Control An eigenmode analysis was performed in AVL to assess the aircraft’s dynamic behavior. This

process used inputs from mass and geometry files which specified the vehicle and operational

parameters, including component surface areas and moments of inertia, operating CL, airspeed, and air

density to calculate the eigenmodes of the system. Each of the conjugate pairs shown in Figure 4.12

characterize a stability mode of the system.



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-20

-15

-10

-50

5

10

15

20

-11 -9 -7 -5 -3 -1 1

Spiral Dutch Roll Short Period Long Period

Figure 4.12 – Vehicle Eigenmodes

4.6.1 Roll Stability

The dutch roll and spiral modes of the aircraft are represented by the blue and green diamonds

respectively, in Figure 4.12. Both of these modes are located in the left half of the complex plane,

indicating that they are stable for all positive time. The dutch roll mode is highly damped, and decays

within 0.60 seconds. The spiral mode is less negative and consequently exhibits a higher settling time,

requiring approximately 2.5 seconds to reach steady state. The highly damped dutch roll mode will

minimize roll-yaw coupling, allowing for more precise heading control, while the positive spiral stability will

provide some self-leveling, reducing the control demands placed on the pilot.

4.6.2 Pitch Stability

The pitch behavior of the aircraft is characterized by a long and short period mode represented by the

pink and red circles in Figure 4.12, respectively. The phugoid mode results from a trade between kineticand potential energy as the aircraft undergoes a series of subtle yet lengthy pitch oscillations. However,

the frequency of the oscillation is sufficiently low, 0.11Hz, so that the long period mode poses no

significant piloting challenges. Since the mode lies in the left half plane, it is stable and will converge over

time. The short period mode lies to the extreme left in the plane and has a consequently short settling

time, 0.43 seconds. This high frequency mode is heavily damped, indicating strong pitch stability.

4.6.3 Center of Gravity

Center of gravity (CG) location, in conjunction with horizontal tail sizing, heavily influences the pitch

dynamics of the aircraft. To promote positive pitch stability, a static margin (SM ) of 15% was selected for

cruise conditions based on a survey of model aircraft SM values. Using Equation 4.3 below and the

aircraft’s neutral point at cruise, a CG position of 4.21 in behind the leading edge of the wing was

selected. It was determined that the payload battery location could be varied sufficiently to account for

changes in CG due to payload configuration.

c

x xSM

cgnp −≡ (Eq. 4.3)



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The aircraft’s neutral point varies, so test cases were run in AVL to assure that the possible CG locations

produced acceptable SM throughout the flight profile. The SM was found to vary from 5.2% at maximum

flight speed to 18% at takeoff, which are within acceptable bounds.

4.6.4 Roll Contro l

Flaps run the full wing span. In order to reduce servo weight, each flap was designed to be

independently actuated as a flaperon. Roll control constraints were dominated by flap sizing required to

take-off in 75 ft. Thus flaperons did not require resizing to satisfy payload roll control requirements.

4.6.5 Pitch Control

Using the resulting horizontal stabilizer area given from Section 4.5.4 as a baseline, the tail geometry

was refined in AVL to satisfy takeoff and cruise trim conditions. Test cases were run in AVL under cruise

conditions (46 ft/s, CL=0.9), constraining elevator deflection to produce a zero pitching moment. The

horizontal stabilizer area and wing decalage were adjusted until the required elevator deflection was less

than 5 deg. This resulted in a lightly loaded tail, which is desirable at cruise as it minimizes the induced

drag contribution from the tail. Additional test cases were run under takeoff conditions (40 ft/s, CL=1.5) to

assure that there was sufficient elevator authority for rotation. Again, elevator deflection was constrained

to produce a positive pitching moment. The horizontal stabilizer area was adjusted such that the tail CL

did not exceed 0.5, above which separation would likely occur.

4.6.6 Landing Gear

The placement and the dimensions of the rear landing gear were driven by the necessary weight

distribution of the aircraft (85% rear and 15% front) and sufficient width and height to prevent a wing-tip

strike. The main gear was placed 15 deg behind the CG and the width adjusted until the wheel contact

patches and CG formed an 80 deg angle, based on historical stability criteria [4]. Due to the relatively

short take-off length, a steerable nose wheel was deemed to be a requirement for ground control.

4.7 Estimated Performance

4.7.1 Aerodynamic Performance

After final sizing of tail surfaces and wing parameters, the MATLAB mission simulation was combined

with lift and drag estimates from a complete Star-CCM+ CFD model. These calculations resulted in the

performance curves shown in Figure 4.13. A maximum L/D of 6.1 was predicted from these models.

Additionally, required thrust at cruise conditions was predicted at 1.90 lbs. and 0.75 lbs for the payload

and delivery missions, respectively. These estimates were compared with wind tunnel and flight testing

of the final vehicle in Section 8.2.



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15 20 25 30 35 40 45 50 55 600

0.5

1

1.5

2

2.5

3

3.5

4

4.5

5

Velocity [fps]

D

r a g [

l b s ]

Level Flight Cruise Drag vs. Velocity

Payload Drag

Payload Stall Speed

Delivery Drag

Delivery Stall Speed

Figure 4.13 – Vehicle Lift and Drag Characteristics

0.1 0.12 0.14 0.16 0.18 0.2 0.22 0.24 0.260.2

0.4

0.6

0.8

1

1.2

1.4

CD

CL

Vehicle Drag Polar

x

4.7.2 Mission Performance

Using the above L/D predictions and updated thrust data from motor testing, the refined preliminary

design was simulated in the primary mission model. Table 4.4 shows the estimated performance from

the model and initial payload testing. System weight was estimated from payload system prototypes,

fuselage prototypes and wing and tail surface area using historical data from past MIT structures and

avionics.

Table 4.4 – Estimated Mission Performance

Delivery Mission Time 3.08 min

# of Delivery Laps 2

Delivery Cruise Speed (max L /D) 28 fps

Cruise CL 0.89

Delivery Energy Consumption 5.35 Wh

Delivery Battery Weight 4.56 oz (6 cell GP1100)

Payload Loading Time (10-50ft) 10-20 sec

Payload Mission Time 2.37 min

Payload Cruise Speed (max L/D) 48 fps

Payload Energy Consumption 8.09 Wh

Payload Battery Weight 9.72 oz (12 cell ELITE 1500)

Predicted System Weight 3.12 lbs.



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5 Detail Design

This section documents the final design specifications for the MIT Team Concrete aircraft.

5.1 Aircraft Dimensional Parameters

Table 5.1 – Aircraft Dimensional Parameters

5.2 Aircraft Structural Characteristics

As shown in Figure 5.1, the wing is the primary loading carrying structure. Landing gear loads are fed

through wing hardpoints to composite A-frame trusses supporting a central spar. This central spar carries

the fabric payload system and supports the tail surfaces. Component structural capabilities are detailedin their respective subsections below.

Figure 5.1 – Aircraft Structure (aeroshell and payload system removed)

General Airframe

Length [t] 4.78

Span [ft] 5.75

Height [ft] 2.28

Wing

Airfoil BAFT2

Root Chord [ft] 0.992

Tip Chord [ft] 0.496

Span [ft] 5.75

Planform Area [ft2] 4.28

Aspect Ratio 7.72Taper Ratio 0.5

LE Sweep [deg] 6.89

Tail

Center Boom Length [in] 46

Boom Diameter [in] 0.575

Horizontal Stabilizer

Airfoil HT14

Root Chord [ft] 0.625

Tip Chord [ft] 0.275

Span [ft] 2

Area [ft2] 0.90

Aspect Ratio 4.44

Arm Length [ft] 2.06

Vertical Stabilizer

Airfoil HT14

Root Chord [ft] 0.75Tip Chord [ft] 0.41

Height [ft] 1

Area [ft2] 0.58

Aspect Ratio 3.44

Arm Length[ft] 2.735



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5.3 Sub-System Design, Capabil ities, and Architecture

5.3.1 Payload Support

In order to minimize system weight, a skeleton structure composed of a central spar and triangular A-

frames were used to support the payload, as opposed to a stressed-skin fuselage.

Central Spar

The central spar is a carbon fiber tube sized to withstand the aerodynamic and payload forces acting

upon the aircraft. In addition to the necessary bending strength required to support the payload, the spar

was also sized for torsional stiffness, due to rudder deflection. It was designed to deflect no more than

1.5 in at a 3g load case and twist no more than 5 deg in a max CL coordinated turn at maximum level

flight speed. Analysis determined that the optimal material was carbon fiber based on its stiffness and

strength to weight ratio. A commercially-available spar that met the specifications was then chosen.

A-frames

It was necessary to find an efficient method for transferring load from the payload system and main

spar to the landing gear and wing. The dimensions of this structure were determined primarily by payload

orientation and spacing. A 0.25 in thick foam-carbon fiber laminate was developed in order to fit between

the payload system pockets. The loads and moments acting on the structure were modeled in SolidWorks

and run through a CosmosWorks Finite-Element analysis. The finite element model included the landing

gear to simulate landing loads and the central spar to simulate the loads associated with the payload and

load transfer between the A-frames. The width of the A-frames was then increased until the structure

could withstand the loading of a full payload, 3g one wheel landing without buckling.

Figure 5.2 – A-Frame Subsystem

5.3.2 Payload Restraint System

Using the lessons learned from the first payload system prototype, additional testing was conducted

to solve an issue with abrasion from the bricks. Strands of Kevlar tow and an extra lamination of Cuben

Fiber in the brick pockets was found to be sufficient for 15-20 loading cycles, more than enough for

competition purposes. A second payload system prototype was then fabricated with the addition of the



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reinforcements and used for subsequent versions. A nylon drawstring and polycarbonate locking cleat

sufficiently restrain all payload configurations while inverted. Kevlar tow is used to tie the payload system

to the A-frames, preventing swaying and excessive movement of the system.

Figure 5.3 – Payload Restraint System

5.3.3 Landing Gear

Front Landing Gear

The main design factors for the front landing gear were steering capability, easy replacement, and

minimum weight. The front landing gear consists of hollow carbon fiber rods with a machined aluminum

single-pivot steering mechanism. The steering servo and push-pull linkage can be seen in Figure 5.4.

Steering capability is necessary in order to maneuver the aircraft in crosswind conditions during takeoff.

The carbon tubes are easily replaced in case of a crash and designed to fail at 15 lbs, 3 times its normal

landing load.

Figure 5.4 – Front and Main Landing Gear

Main Landing Gear

The main design considerations in the development of the main landing gear were easy replacement,

structural strength to survive a 3g landing without permanent deformation, and minimum weight. Further,

the placement and dimensions of the main landing gear were driven by the chosen weight distribution of



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the aircraft (85% rear and 15% front) to allow for proper balance and rotation at takeoff. The main gear

was constructed from bent sheet aluminum with piano wire cross-bracing. It is attached to balsa wood

reinforcements in the wing with nylon bolts.

5.3.4 Wing

Aerodynamic surfaces are of mold-less composite construction. The wing is designed to withstand a

fully loaded 10g turn, obtainable in RC aircraft flight with sudden stick movements or strong wind gusts.

Foam wing cores were cut with a CNC foam cutter and then reinforced with a tapered 2 in root-width

unidirectional carbon fiber spar caps and a single layer of 0.7 oz/sq yard fiberglass oriented at ±45 deg for

torsional stiffness. Balsa hard points and additional bi-directional carbon-fiber skins are located at

attachment points for the payload structure and landing gear.

Figure 5.6 – Completed Wing w ith A-Frames

5.3.5 Empennage

For the tail boom, the aircraft uses the main structural spar mentioned earlier. The carbon fiber tube

is epoxy-bonded t

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