The Balsa bullet: A high speed, low-cost general aviation aircraft for Aeroworld

Appendices A-1 Appendix A: Deliverable Items B-1 Appendix B: Critical Data Summary C-1 Appendix C: Aircraft Data Base D-1 Appendix D: Performance Program E-1 Appendix E: Aerodynamics, Stability, and Control Program F-1 Appendix F: Manufacturing Plan 1.0 EXECUTIVE SUMMARY Long Shot Aeronautics has designed the first general aviation aircraft to service Aeroworld. In accordance with the mission definition provided by AE 441, INC. management, Long Shot Aeronautics parent company, The Balsa Bullet, shown in Figure 1-1, is a high speed low cost six passenger general aviation aircraft. It will cruise at a speed of 55 ft/s with a maximum speed of 75 ft/s for distances in excess of 27000 feet. This range and speed combination provide The Balsa Bullet with the capability to service any two existing airports in Aeroworld in an efficient and timely manner.

Overall, three major design drivers have been identified by the design team. The first is to provide a low cost airplane to the Aeroworld market.

Maintaining the low cost objective will not simply meet the mission objective as defined by AE 441, INC. management but will also make the Bullet an economically viable option for a wide number of consumers. The Balsa Bullet has a total manufacturing cost of $1000 with a price to the consumer of only $2562. The second major driver is high speed performance. Once again this driver exists not only to meet the mission objective given Long Shot Aeronautics but it provides a desirable feature to the consumer, pride in owning the fastest aircraft in Aeroworld. The third design driver identified is the capability to service any runway in Aeroworld necessitating the ability to takeoff within 28 ft, the length of the shortest runways in Aeroworld. These design drivers provide three great reasons for the general public to purchase a Bullet.

The propulsion system consists of a Zinger 12-8 propeller mounted on a powerful Astro-15 motor located in the nose of the aircraft for symmetric thrust as well as weight balance. The motor is powered by thirteen 1000 1-1 Figure 1 - 1: Three View Sketch of The Balsa Bullet Sw = 6.33 ft 2 Sv = .50 ft 2 cw = .96 ft cv = .41 ft Sh = 1.50 ft 2 length = 3.54 ft bh = 2.74 ft Zin_;er 12 - 8 propeller Scale: I inch = 1.74 feet I t I i

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'l I f 1-2 Fi_m_ 1-2: In_ Configuration wing spars avionics batteries passengers motor speed controller servos batteries 4 channel receiver speed controller

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_IirC__ _ batteries wing spar motor /jt_wi ng_spar servos passengers 4 channel receiver and avionics batteries I-3 milli-Amp hour Panasonic batteries located in the wing carry-through structure. The propulsion system is powerful enough to provide a cruise speed of 55 ft/s and a takeoff distance of 21 feet which will allow The Balsa Bullet to serve any airport in Aeroworld. The maximum takeoff weight is 4.60 pounds.

The aircraft configuration is a low rectangular wing monoplane with five degree wing dihedral. A design constraint of executing a level 60 ft radius turn at a speed less than 30 ft/s was placed on the aircraft. In relation to the cruise and maximum speed, this turn speed is low. As a result, a tradeoff exists between the large wing desired for low speed flight and the small wing desired for high speed flight. Selection of the wing size and characteristics had the most significant influence on the final design. In order to achieve the highest lift with the smallest wing possible, the FX-63-137 airfoil was chosen. The employs plain flaps extending twenty percent of the 11.52 inch chord and half the 6.33 foot span.

A rectangular fuselage provides each passenger with a spacious 18 in 3 seating and generous room for 9 in 3 of baggage. The rectangular fuselage was chosen over a more aerodynamic circular cross section because it requires less time to build thereby yielding a lower manufacturing cost consistent with the designs major objectives. The fuselage has an access hatch on the top for propulsion removal or installation in under twenty minutes.

The landing gear of The Balsa Bullet is exceptionally long as a result of the three inch grass outdoor takeoff requirement. The tricycle configuration was chosen to provide increased ground stability while providing fuselage grass clearance. The nosewheel landing gear is 8.0 inches long and the main gear are 7.8 inches long to provide propeller clearance. The main gear is slightly shorter to provide an aircraft take-off angle of attack of 2.25 degrees for 1-4

increased lift prior to rotation without further complicating wing attachment

with an incidence angle.

The center of gravity of the Bullet at maximum takeoff weight is

located at 27% of the mean aerodynamic chord. The neutral point of the

aircraft is located at 49.2%of the mean aerodynamic chord, producing a static

margin of 22.3%of the mean aerodynamic chord.

The control systems of the Bullet consist of a rudder which is fifty

percent of the 0.5 square foot vertical tail and an elevator which is twenty

percent of the 1.5 square foot horizontal stabilizer. Roll control is provided by

rudder deflection and wing dihedral. A steerable nosewheel is also provided

for improved ground handling. Unlike previous Aeroworld designs, The

Balsa Bullet employs airfoil sections in the horizontal and vertical tails. The use of NACA 0009 airfoils reduces the drag while increasing aircraft performance and resistance to twist.

The critical technologies identified by the design team include flap effectiveness, mating the wing with the fuselage, and team manufacturing experience. Flaps are a concern due to uncertainty in their effectiveness on past designs. On some past aircraft, flaps appeared to be ineffective at best and sometimes a detriment while on other aircraft they seemed to be beneficial. It is difficult to ascertain whether airplanes employing flaps that proved detrimental or ineffective resulted from poor manufacturing of the wing and flap or if earlier designs failed to consider the drag increase associated with flaps and were unable to overcome it with their propulsion system. The wing and fuselage mating are a concern in two areas: attachment to the fuselage floor due to high stress concentrations, and at the edge of the fuselage where a tight seal must be made at the opening in the monokote where the wing carry-through enters the fuselage.

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A critical technology which cannot be overlooked is the lack of

experience of the design team in manufacturing. The lack of experience may

have a negative impact on the final product and hurt the prototype's

performance as a result of poor manufacturing rather than poor engineering.

The Balsa Bullet is not without it's weaknesses. Its predicted range will not allow it to fly directly from any airport in Aeroworld to any other airport including a one minute loiter and diversion to the nearest alternate airport.

Though the Bullet was not designed to travel long distances non-stop, a competing aircraft of similar cost possessing a long range capability would be more attractive to a consumer. Also, the design is not aesthetically pleasing as a result of the square fuselage, wing, and empennage. These drawbacks may discourage some clientele from purchasing The Balsa Bullet.

Although The Balsa Bullet has these drawbacks, its economic appeal and high speed capability far outweighs them. The Balsa Bullet's affordability to a relatively large number of consumers will make it a success in Aeroworld, particularly as the first entry into the general aviation market.

1-6 Figure 1 - 2: Summary of The Balsa Bullet Specifications Empennage Aerodynamics Horizontal Tail Area 2.958 sq ft 6.33 sq ft Wing Area NACA 0009 Horizontal Tail Airfoil 7.2 Aspect Ratio Horizontal Tail Incidence -2.25 deg 6.75sq ft Span Angle 0.937 ft Elevator Area Mean Chord 0.75 sq ft Max Elevator Deflection 30 deg Taper Ratio Vertical Tall Area 0.5 sq ft Sweep 0 dei_ NACA 0009 Vertical Tail Airfoil Dihedral 5 deg Rudder Area FX63-137 0.25 sq ft Wing Airfoil Section Max Rudder Deflection 30 deg Wing Incidence Angle 0 deg 0.633 sq ff Flap Area Max Flap Deflection 20 deg 0.0288 CDo Performance Propulsion Astro 15 21.3 ft Take-OffDistance Engine 13 1.2V 1000mall 20.45 ft/s Batteries Velocity (min) 84.6 ft/s Zinger 12-8 Propeller Velocity (max) 0.528 lb 13.92 min Thrust (cruise) Endurance (max) 28692.2 ft Thrust (max) 3.12 lb Range (max) Propeller RPM (cruise) 4793.8 Economics Structure 4.6 lb Direct Operationg Cost $4.18 Max Weight 40 in Cost per 1000 feet $0.26 Fuselage Length Total Aircraft Cost $2562.11 3.5 in Fuselage Width 3.5 in Fuselage Height 1-7 2.0 DESIGN REQUIREMENTS AND OBJECTIVES AE 441, INC. management has decided to market the first general aviation aircraft to service Aeroworld. Currently, Aeroworld is serviced only by commercial transports thus leaving an entire market untapped. It has been determined that the first general aviation aircraft introduced to Aeroworld will be a low-cost, high speed, six passenger, electrically powered flight vehicle that will provide benign handling qualities without sacrificing performance. Lastly, the aircraft must be able to be mass produced.

2.1 Target Market A map of Aeroworld, the Bullet's intended market, appears in Figure 2-1. As one can see, the majority of runways are 40 ft long but both City C and City O have runways which are only 28 ft long. Long Shot Aeronautics wishes to serve all cities in Aeroworld as both a convenience to our customer as well as maximizing availability in the market. Thus a 28 ft takeoff distance objective is imposed on the design by the design team.

Since a commercial transport fleet exists and it is not the nature of general aviation aircraft to fly great distances non-stop, the Bullet is targeted for a range of 16000 ft including diversion to the nearest alternate airport and a one minute loiter. By not trying to compete against the lower cost of commercial transports over great distances, the Bullet will require lower capacity batteries. Lower capacity batteries are cheaper to purchase than high capacity batteries and are lighter as well which provides a weight savings that translates into an additional cost savings.

2-1 Figure 2 - 1: Aeroworld Aeroworld Airport Information Below is a map of Aeroworld as well as a table of the latitude and longitude of each of the cities. Each latitude and longitude increment represents 500 ft.

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./. -20 • LONGITUDE Runway City Longitude Latitude Length Factor A -21 6 1 B -15 12 0.8 C -10 -5 0.7 D -1 -10 1 E 9 -1 1 F -4 10 1 G -5 17 I H -1 12 1 I 8 7 1 J 5 15 I K 9 17 I L 1 20 15 M 1 20 5 N 1 24 10 O 20 -9 0.7 2-2 Figure 2-2 shows the numbers of flights for a given range between the airports of Aeroworld. With a design range of 16000 ft, The Balsa Bullet will serve approximately 56% of all possible flights. This percentage increases to 84% without loiter and diversion. It is important to note that all cities in Aeroworld are able to be served with a single stop.

For the market targeted by Long Shot Aeronautics, three primary design drivers have been identified: high speed to make the aircraft performance attractive to the consumer; low cost so a large number of consumers can purchase the Bullet ; and able to takeoff within 28 ft to enable the airplane to service every city in Aeroworld.

°_ © --_ 10 Z r'M '¢" & & _ ,4 ,6 & & & Range in Thosands of Feet Figure 2 - 2: Number of Flights for a Given Range 2-3

2.2 Requirements

The following are the performance and manufacturing requirements

and objectives for the proposed aircraft to operate successfully in Aeroworld

as determined by AE441, INC. upper management and meet the constraints

imposed by the indoor and outdoor flight test environments.

2.2.1 Performance

• Execute a level 60 ft radius turn at a speed less than 30 ft/s

• Fly a closed figure eight courseless than 40 yds x 100yds

without exceeding an altitude of 25 ft

• Clear a 50 ft obstacle within 200 ft on open course with 3 inch

grass rough field characteristics

• Maximum take-off distance of 40 ft indoors, 60 ft outdoors

• Range capability to service any two existing airports with

stops for refueling

2.2.2 Manufacturing

• 6 passengercapability

• 4 in3 baggageper passenger

• Battery placement in wing carry-through structure

• Minimum propeller clearance of 3 inches

• Propulsion system installation in under 20 minutes

• $290.00maximum limit on raw materials

• Maximum of 4 servos for a 4 channel transmitter

• 2 week maximum construction time

• Avionics crash survivability

• Must meet all pertinent FAA and FCC regulations

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2.3 Objectives The following are the performance and manufacturing objectives of Long Shot Aeronautics for the production of The Balsa Bullet. These are the characteristics deemed necessary and attainable to produce a high quality aircraft to the consumer.

2.3.1 Performance • Minimum range of 16000 ft including loiter and diversion • Maximum velocity: at least 75 ft/s • Cruise velocity: at least 55 ft/s • Maximum takeoff distance indoors of 25 ft 2.3.2 Manufacturing • Maximum manufacturing time of 90 hrs • Weight estimate: 4.60 lb max • Manufacturing cost: $1600.00 max • Minimum load factor of +2.0/-1.0 • Crash survivability of 12.0 lb • 8 in 3 per passenger seating 2.4 Exceptions From Original DR&O There are no exceptions to the DR&O submitted 25 January 1994.

2-5 3.0 CONCEPT SELECTION 3.1 Preliminary Concept Proposals In order to produce a design to meet the requirements set forth by AE 441, INC. management, each of the six design team members submitted individual preliminary concept proposals. Electric engines are considered state-of-the-art in Aeroworld and appear in each concept. Only single engine concepts were considered in order to avoid the significant cost of purchasing a second engine in keeping with the design team's low cost objective of not exceeding $1600.00 in purchasing and manufacturing costs. From these six proposals the design team has the option to select one of the designs or attributes from each of the six designs to produce the final product. An overview of each proposal will be presented followed by the final design selection and rationale. Traits common to all the concepts will first be explained, however it should be noted that the assertions made are qualitative in nature due to the time constraints imposed in the design.

Quantitative analysis was reserved for the final design selected.

In each preliminary design concept a rectangular fuselage was proposed. This results from the desire to produce a reliable low cost aircraft.

A circular fuselage, while more aesthetic and aerodynamic, would be much more difficult to manufacture. The difficulty lies in making a circular cut across the grain of soft woods chosen for their low weight without splintering the piece being tooled. Additionally, there would be an increased tooling cost for circular cuts as opposed to straight cuts used in rectangular fuselage manufacture. Bending the wood used in longerons is also time consuming and requires precision to achieve a consistent desired circular fuselage shape.

This difficulty in manufacturing the aircraft would lead to higher prices as a 3-1

result of the increased manufacturing time in turn damaging the design

team's low cost objective. Additionally, Long Shot Aeronautics is compelled

to build as simple a design as possible due to a lack of manufacturing experience.

Every preliminary design also suggested the use of tricycle landing gear with a steerable nose wheel. Tricycle landing gear are attractive for three major reasons. First, the possibility of ground looping which exists for a tail dragger configuration is not present with the tricycle gear. Secondly, a tailwheel would not raise the tail end of the fuselage out of the three inch grass increasing the takeoff roll. Third, a tricycle landing gear configuration also maintains a forward center of gravity position thus allowing for a smaller horizontal tail for control as a result of the increased moment arm.

A final attribute common to all six proposals entailed use of a rectangular wing. Taper would require different size ribs along the span whereas a rectangular wing uses ribs of a single size. All of the ribs can be cut at one time from a template for the rectangular wing thus lowering manufacturing time and overall cost. The Bullet does not operate at flight speeds which make sweeping the wing beneficial. Sweep would add a second angle in addition to the dihedral that must be accounted for in attaching the wing to the fuselage. This second angle is detrimental as the design would be complicated and probably require additional material to successfully attach the wing to the fuselage thereby increasing the cost.

3.1.1 Concept A Concept A (see Figure 3-1a) is a high wing monoplane design with rectangular fuselage and a V-tail. The design employs tricycle landing gear and a rectangular wing with flaps. The high wing was chosen to provide roll 3-2

stability and simplify the manufacturing process by eliminating wing dihedral

and allowing the wing to be attached to the upper surface of the fuselage by a

simple bracket. By having a removable wing, the interior of the plane

becomes more accessiblewhile allowing the wing to be manufactured as a

part separate from the fuselage construction. A major disadvantage of this

design results from the requirement for battery placement within the wing

carry-through structure. In doing so, a platform would need to be constructed

to support the batteries. In addition to increased material costs, the center of

gravity would be raised farther above the ground thereby increasing the

possibility of tipping over on the necessarily long landing gear.

The V-tail was chosen as a way to reduce interference effects from the

wing on the tail surface. Wing interference is important due to the short fuselage associated with a general aviation aircraft. The V-tail is complicated from a controls perspective as the system must allow for both synchronous as well as differential actuation of the control surface. Attaching a V-tail to a rectangular cross section would require an increase in structure over a simple cruciform tail due to the angle of attachment and loading experienced by the tail. Difficulty in construction, coordination of the control surface actuation, and a possible increase in weight and cost associated with the V-tail made this option unattractive.

Flaps are proposed as a way to provide the necessary lift for takeoff from all Aeroworld runways. The major advantage that results is a smaller wing. A smaller wing is more efficient at higher speeds than a lower wing.

The disadvantage of flaps is a wing that is more difficult to construct than a wing without control surfaces. There is also a production cost increase associated with the purchase of a servo to control the flaps, additional weight from structure necessary to support the flaps, and increased time to 3-3

manufacture a wing with flaps over a wing without control surfaces. This

cost can be offset in the long run through a decreasein operating costs

associatedwith a more efficient cruise at a higher lift-to-drag ratio. The

weight increase for additional structure for the flaps is offset through the

decreasedsize of the wing when flaps are used.

3.1.2 Concept B

Concept B (Figure 3-1b) is a high rectangular wing monoplane with

tricycle landing gear, cruciform tail, ailerons, and side by side passenger

seating. The high wing was chosen for the same reasons as those listed in

Concept A and carries with it the same disadvantages.

Side by side passengerseating was chosento easepassengeraccessto

seating and baggage areas. In doing so the fuselage has a greater frontal area

than single file seating and a much greater profile drag. Additionally, the

shorter fuselage associatedwith side by side seating will increase the

horizontal tail area for control purposes as well as increase wing interference

on the tail.

Ailerons were suggested as a way to improve the control of the aircraft

in turns. The roll control provided by ailerons also minimizes the skidding

and slipping sensations experienced by the passengers and pilot associated with maneuvering. The drawback of using ailerons instead of flaps is increased wing area leading to less efficient high speed characteristics, particularly in cruise, which will increase operating costs. Ailerons also incur a weight penalty because unlike flaps, there is not a decrease in wing size to offset the additional structure needed.

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3.1.3Concept C

Concept C (Figure 3-1c) is a low wing monoplane with tricycle landing

gear, cruciform tail, tapered rear fuselage, and a canopy. The low rectangular

wing employs dihedral to provide roll stability with ailerons for roll control.

The low wing was chosen for improved aesthetic qualities as well as allowing

battery placement lower to the ground. As proposed, the low dihedral wing

will be more difficult to attach to the fuselage than the high wing designs for

several reasons. The low wing is permanently mounted at an angle to the

fuselage floor. This angle requires materials strong enough to sustain the

concentrated loads at the point of attachment. The mating of the wing and

fuselage at the root is difficult because a hole in the fuselage monokote must

be made for the wing carry-through structure. This interface must be

effectively sealed in order to not jeopardize aerodynamic integrity. These

aspects of the design illustrate that unlike the high wing which can be built

separate from the fuselage, the low wing must be built in close association

with the fuselage.

The aft body is tapered for improved aerodynamics. Aft body tapering,

although not extremely difficult, is more difficult than running a continuous

straight beam down the length of the fuselage due to the need for an interface

from which the tapering can begin. An increase in the number of joints

increasesthe weight due to use of more glue. Taper may also decrease

structural integrity barring increased structure due to the angled interfaces as

well as complicate load path determination.

The use of a canopy on Concept C is a feature not found on previous

Aeroworld designs. Previous designs internalized the pilots and passengers

without really providing a way for the pilots to see forward of the aircraft

without engine obstruction. The canopy could also be used as the interior

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Figure 3 - 1: Concepts A, B, and C r ; i ,J iI A ,4 °_ 3-6 access panel to meet the propulsion system removal requirement and provide interior access. An obvious drawback is the increased drag associated with a canopy over internalization of the pilot and passengers within the fuselage.

3.1.4 Concepts D, E, and F These concepts, appearing in Figure 3-2, incorporate attributes previously mentioned in Concepts A, B, or C. As such they will have the same advantages and disadvantages previously noted.

Concept D is a high wing monoplane with cruciform tail, flaps, and canopy. Concept E is a low wing monoplane with side by side passenger seating, a canted forward fuselage to allow the pilots to see over the engine, fore and aft taper, and a cruciform tail. Concept F is a high wing monoplane employing a V-tail and midspan ailerons.

A summary of the major aircraft characteristics with their associated strengths and weaknesses appears in Table 3-1.

Weaknesses Feature Strengths High wing • inherent roll stability • requires additional floor for •easier to manufacture batteries in carry-through • raises aircraft c._;.

• meets battery placement • requires dihedral to provide Low wing roll stability requirement • more difficult to build • lowers c.g.

• lose ailerons Flaps • decrease wing size •improve cruise and takeoff • poor performance on similar win S in earlier aircraft performance Ailerons • increase roll control and • lose flaps and associated benefits overall handlin_ qualities • provides forward view Canopy •increases drag significantly V-tail •reduces wing interference on •increased weight and tail difficult to build • easy to build therefore Square fuselage • not as aerodynamic as a cheaper circular fuselage Table 3-1: Strengths and Weaknesses for Various Concept Attributes 3-7 Figure 3 - 2: Concepts D, E, and F ¢j O 6_ I ,w.G !

3-8 3.2 Final Design Selection The final design (Figure 3-3) chosen maintained the traits common to all six design proposals: tricycle landing gear, rectangular shaped fuselage, and a rectangular wing for the advantages stated earlier. A cruciform tail was chosen due to the manufacturing difficulties associated with a V-tail and because a large fleet of aircraft using the cruciform design are available for reference. Due to the battery placement requirement and the necessary use of unusually long landing gear for the outdoor flight test requirement, a low wing aircraft was chosen. This decision increased the final design's stability through a lower center of gravity and removed the need for additional structure other than a floor to support the batteries thereby reducing cost.

As the design analysis began, the configuration included ailerons instead of flaps. However, initial aerodynamic and performance analyses indicated the need for a smaller wing during cruise to achieve the necessary speed objectives and avoid large negative angles of attack to maintain a level cruise altitude. Without flaps, the large wing required to meet the stated takeoff distance of 25 ft would require a significant negative angle of attack to prevent the aircraft from climbing at the cruise speed assuming the increased drag associated with the wing and fuselage at this attitude could even enable the Bullet to reach its cruise velocity. The decrease in cruise efficiency would also impact negatively on the operating cost of the airplane. Flaps will also help meet the short field takeoff objective of the group. Given these factors the decision was made to switch to flaps from ailerons.

A canopy was not chosen due to the inherent drag penalty and negative effect on aircraft high speed performance, a major design driver. In light of the number of successful flights made in Aeroworld in aircraft without canopies, this decision seems to have little, if any sacrifice.

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A feature which does not appear on the preliminary proposals but does

appear on the final design is the use of an airfoil section for both the

horizontal and vertical tail. By using an airfoil, the design experiences less

drag than the flat plate sections appearing in the proposals. This drag decrease

provides the Bullet with an increase in performance. Additionally, an airfoil

is less subject to twist than a flat plate, thereby assuring a consistent response from the aircraft in flight.

3- 10 Figure 3 - 3: The Balsa Bullet 3-11 4.0 AERODYNAMICS The aerodynamic design of The Bullet, especially the wing design, was the most difficult task in the design process. The high-speed(55 feet/second) versus short takeoff distance (25 feet) conflict was the primary driving factor in the design. Although this conflict was the primary design driver, cost and ease of manufacturing also played a significant role in the aerodynamic design. These drivers led to the major dilemma of whether or not to use flaps. The decision to implement flaps was made rather late in the design process to increase efficiency and decrease wing area.

It should also be noted that Reynolds number effects are a critical aerodynamic issue since The Bullet flies in a relatively low Reynolds number regime (Re = 350,000 for a cruise speed of 55 ft/s and chord length of .94 ft ).

4.1 Airfoil Selection The selection of an airfoil section for The Bullet was driven primarily by two main factors: Design cruise speed goal of 55 feet/second Takeoff distance of 25 feet First, good aerodynamic performance is paramount not only for high speeds, but also for lower costs. The airfoil must exhibit low drag characteristics. Also, high lift characteristics are essential to attain takeoff requirements. Second, the geometry of the airfoil is crucial to the design of The Bullet. The airfoil must be at least an 1.25 inches thick to allow the batteries to be placed in the wing.

Although this added thickness is a weight penalty, the structural resistance to longitudinal wing twist is reduced. Finally, because ease of manufacture is an important concern as well, airfoil geometry again becomes relevant.

4-1 In the low Reynolds number regime in which The Bullet will be operating, there are several airfoil sections which merit consideration. Using data from Reference 13, Table 4-1 lists several of these airfoil options.

Airfoil Max C1 % % Camber Cd @ 0 deg Thickness Re = 300,000 Aquila 1.3 @ 12 deg 9.38 4.05 .03 Clark Y 1.2 @ 10 deg 11.72 3.55 .011 WB 140-35 1.15 @ 10 deg 13.92 3.7 .01 FX-63 137 1.6 @ 12 deg 13.7 5.94 .01 TABLE 4-1: AIRFOIL DATA In order to select an airfoil that meets the mission requirements some type of trade-off must occur between the strengths and weaknesses of particular airfoils. In order to meet the short takeoff distance without employing flaps, a high C 1 is needed. Keeping in mind the requirement that the airfoil be at least 1.25 inches thick to place the batteries in the wing carry-through structure, thickness becomes important. The WB 140-35 was eliminated from consideration due to its low C 1, even though it was the thickest of the airfoils. The Aquila has a significantly less thickness ratio than the other airfoils, which means the chord would have to be 20% larger to fit the batteries in the wing carry-through structure. For the rectangular planform of The Bullet, the net result would be a lower aspect ratio and consequently decreased aerodynamic performance. In order to achieve high speeds, drag must be minimal. Therefore, the L/D ratio for the airfoil must be considered. The L/D for the Clark Y is about 40 compared to 78 for the FX-63 137 at Re = 300,000.

The trade-off in choosing the increased aerodynamic performance of the FX-63 137 is a large moment coefficient (Cmo = -.24) and increased manufacturing difficulties. This large Cmo would effect trim characteristics of the airplane. The FX-63 137 was finally selected over the Clark as the airfoil 4-2

section solely because the need for high lift was a higher priority than ease of construction. It should be noted that this decision was made prior to the decision to use flaps. Therefore, a different airfoil may have been chosen if flaps were considered.

The FX 63-137 lift curve is shown in Figure 4-1. The maximum C1 is 1.6 and the lift curve slope is approximately .1/deg.

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Re = 200,000

Cmo = -0.24

Cd = .01 @ 0 degrees

O Clmax U') O

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tO O O tD I c_ Figure 4-1:FX63-137 Airfoil Lift Curve 4-4

4.2Wing Design

4.2.1General

The wing design of The Bullet was the primary driver in the whole

aircraft design. The characteristics of the wing played a crucial role in every aspect of the design decision making process. The primary goal of the wing design was to obtain the smallest wing area possible while still meeting all design requirements and objectives. Because of the requirement that two wings be fabricated for testing purposes, a rectangular planform was chosen with little hesitation due to its simplicity. Simple geometry also translates into less manufacturing time and ultimately less cost.

4.2.2 Wing Sizing The main requirements that determined the sizing of the wing were cruise performance takeoff performance turn performance wing loading AE 441, Inc., requires that the plane be able to execute a level 60 foot radius turn at a speed less than 30 feet/second. A takeoff distance of no greater than 25 feet was also an objective so that all airports in Aeroworld could be serviced by The Bullet. The wing loading must also be taken into consideration for the RPV's of Aeroworld. Management recommended that the wing loading should not exceed 12 ounces/foot, to ensure takeoff capabilities and structural soundness within the design.

In order to optimize the wing area, a computer program was written which determined the wing area using the aforementioned factors. A copy of the program is attached as Appendix G. The wing sizing process is a highly iterative 4-5

processbased on several factors. Therefore, several assumptions were made to

determine the wing area. First, and most importantly, a good weight estimate was required. As the design process evolved, the best weight estimate was 4.6 lb.

The velocity at takeoff and during turns must also be known. Finally, the value of C1 is necessary to size the wing. CL for the aircraft can be found by modifying the airfoil lift curve slope to account for three-dimensional down wash effects if the aircraft efficiency factor and angle of attack are known. The efficiency factor, e was assumed to be .8. To assure the passengers a comfortable ride, an angle of attack of 5 degrees was used during turns, while the angle of attack during takeoff was assumed to be 10 degrees without flaps to avoid stall.

Using a weight of 4.6 lb., the program solves for the wing area at both takeoff and during turns while constraining the wing loading to no more than 12 ounces/foot. For the turn performance case, the load factor for a 60 foot level turn at 25 feet/second was found using the relation: Using this load factor the wing area can be found for turns using: nW S= .$CLpV 2 Using a load factor n = 1 for takeoff along with the CL from the methods described above, the wing area can be found for takeoff assuming Vtakeoff = 25 feet/second. The larger of the two areas determines the wing size. Prior to using flaps, the wing area converged at 7.2 sq.ft., while it converged to 6.3 sq.ft.

using flaps in the design.

The aspect ratio of the wing was then found by imposing the restriction on the airfoil that the batteries be placed in the carry-through structure. Since the airfoil thickness is a function of chord, the chord for The Bullet was fixed at .94 feet. Knowing the area and chord, the span was found to be 6.75 feet, which 4-6

results in an aspect ratio of 7.2. The aspect ratio is an important design

parameter since it determines the efficiency of the wing. As the aspect ratio increases better aerodynamic performance is expected, but the penalty paid results in a weaker wing structure. The aspect ratio of 7.2 imposed on The Bullet will not hinder the aerodynamic or structural performance of the design.

It should be noted that the CLmax for the aircraft was 1.3 without flaps.

The value for CL max was obtained by multiplying the value of the aircraft lift slope (.077/deg) times the angle of attack of the airfoil, which is the stall angle of the_(12 deg) minus the zero lift angle of the _(-6 deg). Knowing the value of CL max to be 1.3 the CL at takeoff was assumed to be 1.2, which is slightly less than C L max. Knowing the lift coefficient at takeoff and the aircraft lift curve slope, the necessary angle of attack at takeoff can be determined. For The Bullet, the following relationship holds if the ground is the reference line.(i.e.

the relative wind is parallel to the ground).

C L = Ct_ (i_,._ - (z_ - i,) where iground is the angle the fuselage makes with the ground and iw is the angle at which the wing is mounted to the fuselage. In order to avoid construction difficulties that would arise by mounting a low-wing structure to the fuselage at an angle, iw was kept at 0. Because the zero-lift angle is constant for the wing at -6 degrees, the angle at which the fuselage sits with respect to the ground becomes important. Since this is a maximum finite angle for a given fuselage length and landing gear length, the necessary CLmax to reach a takeoff distance of 25 feet could not be achieved without mounting the wing at an angle to the fuselage. Rather than mounting the wing at an angle, the decision to use flaps was implemented.

4-7 4.2.3 Flaps The decision to use flaps was not one of necessity, but rather one of efficiency. As mentioned previously, the decision was made rather late in the design phase for several reasons. Because two of the main goals of the aircraft are to achieve high-cruising speeds and short take off distances, the lift coefficient must be as high as possible during takeoff. This can be achieved by using flaps or highly cambered airfoils. However, when highly cambered airfoils are employed, the wing is less efficient in cruise. Also, since takeoff is the primary controller of wing area, the addition of flaps can reduce wing area and increase efficiency at cruise. Using an aircraft weight of 4.6 lb., Table 4-2 illustrates the benefits when flaps are employed during takeoff.

Weight = 4.6 lb. [ Wing Area L/D @ cruise V = 55 ft/s Bullet w/flaps (.2c ,.5b) 6.3 sq.ft. 8.5 Bullet No Flaps 1 7.2 sq.ft. 3.8 (Flaps @ Takeoff only) TABLE 4-2 BENEFITS OF FLAPS Of course, the penalty for flaps is a larger drag at takeoff, increased production costs, and manufacturing difficulty. Also, flaps were a design risk because their effectiveness has been uncertain on previous RPV's in Aeroworld due to increased drag. However, the decision to use flaps was made to resolve the conflicting mission requirements of high cruise speeds and short takeoff distances and decrease wing area.

The sizing of the flaps was determined by takeoff performance as well.

The takeoff distance for various flap chord sizes and deflections is illustrated in Figure 4-3.

4-8 m 10 degree flap deflection • 20 degree flap deflection A 30 degree flap deflection ..l=J .,,-4 ¢3 © I [...,

S = 6.'. sq.ft. "_

W = 4 6 lbs

20- Vtakeq _ff= 25 ft, 5 l I I I I ! ! I I | ! ! I ! I ! I ! I I 0 0.4 0.5 0.1 0.2 0.3 Ratio of Flap Span to Total Span Figure 4-3: Effect of Flaps on Take-Off Distance 4-9

A flap size of at least 20% chord was desired for manufacturing reasons and a

maximum deflection of 30 degrees was allowed. Takeoff and manufacturing considerations set the flap size at 20% chord and 50% span with a maximum deflection of twenty degrees to achieve the takeoff distance of 25 ft.

With the size of the flaps set at 20% chord and 50% span, their effect on the airfoil was determined for a maximum deflection of 20 degrees. Using the methods presented in Reference #. The flaps were found to increase the C1 of the airfoil by .2. However, the flaps also increase C d by a factor of .02. The use of flaps does not effect the lift curve slope of the aircraft, but they do shift the curve up and decrease the stall angle. The final aircraft lift curve slope both with and without flaps is shown in Figure 4-5. These results were used in conjunction with the computer program in Appendix G to determine the final wing sizing and aircraft aerodynamic characteristics.

1.6 / w/flaps @ 9 deg i Stag @ 1_1deg 1.2

1.4 _" r"' "

, .jd.; _ r g,- .__

/ /

/ _,_c,

= M J

"i 0.4 CI flaps@ (20deg) 0.2

o ?._r

-0.2 -8 -6-4 -2 0 2 4 6 8 10 12 Angle of Attack (deg) Figure 4-5 Aircraft Lift Curve 4-10 4.2.4 Dihedral The decision to use flaps on a low-wing monoplane meant that ailerons were eliminated from the design, since all four available servos were now in use.

The low wing design was chosen to avoid extra structural and cost penalties associated with placing the batteries in a high wing. The consequence of these decisions is that dihedral must be incorporated into the design to provide the necessary roll stability. As shown in Section 7-6, the necessary wing dihedral was 5 degrees.

4.2.5 Load Distribution With the wing characteristics fixed, the next important parameter is the load distribution. The load distribution is needed during takeoff and at cruise conditions for structural considerations. Takeoff becomes especially important since this is when the wing will experience its greatest loads.

--N7

0.9 0.8

"q \\

0.7 b, (9 0.6 • i,,,_ 0.5 O -I-- L/Lr takeoff G; 0.4 (Flaps down) 0.3 0.2 + L/Lr cruise

\

0.1 (No Flaps)

\

0- 0.5 1 1.5 2 2.5 3 3.5 Spanwise Location (ft) FIGURE 4-6: LOAD DISTRIBUTION 4-11

The load distribution was analyzed using a lifting-line code written in

Aerodynamics 350. The results show what one would expect. The load distribution is approximately elliptical as shown by the curve fits. It should be noted that the load distribution changes greatly when flaps are deployed. The change in CL is .2.

4.2.6 Final Wing Characteristics The final wing characteristics of The Bullet are shown in Table 4-4.

Wing Area 6.3 sq.ft Span 6.75 ft Aspect Ratio 7.2 Taper Ratio 1.0 Dihedral 5 deg Chord .94 ft ew .88 Wing Loading 11.6 oz/sq.ft.

TABLE 44: FINAL WING CHARACTERISTICS 4-12 4.3 Drag Prediction The estimation of the drag coefficient is a difficult and challenging task even for the simplest of aircraft such as the RPV's of Aeroworld. The standard method used in determining aircraft drag prediction requires the drag to be split up into the parasite drag and induced drag. The governing equation can be found in most aerodynamic textbooks.

CD =CD_ + C L _A Re Although this equation seems simple, the real hurdle is predicting both the parasite and induced drag contribution for each aircraft component. Numerous methods are available with each one having its strengths and weaknesses. In order to obtain the most accurate drag prediction possible, three different methods were applied to our configuration.

4.3.1 Nelson's Method An initial calculation for C_was performed using the component buildup method shown in Reference 11.

The reference area used in the analysis was the wing area S = 6.3 sq. ft. The results are presented in Table 4-4.

Component CI_ A_ (sq.i_) (Source of A_) CDo 0.007 6.3 Wing 0.007 Swing 0.11 0.0851 Fuselage 0.0015 Fuselage Max frontal area Horizontal tail 0.008 1.5 Hor. Tail Area 0.0019 Vertical tail 0.008 0.5 Vert. Tail Area 0.0006 0.014 6.3 Landing Gear 0.014 Swing Interference 15% Total CDo = 0.0267 TABLE 4-4: DRAG BREAKDOWN : NELSON METHOD 4- 13

4.3.2 Jensen's Method

Since Nelson's method applies to real world aircraft, it was necessary to find an alternative method which could account for the low Reynolds number flight regime of Aeroworld planes. Using Daniel T. Jensen's A Drag Prediction Methodology for Low Reynolds Number Flight Vehicles, a more detailed estimate of the drag can be found.

The results are presented in Table 4-5. The CDo from the wing was found using data in Jensen's thesis for the FX-63-137 wing section at a Re of 300,000.

Cf_ FF_ Component Swetx (sq.i%) CDo 4.06 0.0021 0.003 1.0832 Fuselage Body 0.0013 Horizontal tail 0.0032 0.829 1.5 0.0005 Vertical tail 0.0035 0.829 0.5 0.0118 Wing ( using FX-63 137@ Re=300000) 0.014 Landing Gear (From Previous Method) 0.0024 Interference 15% = .0321 I TOTAL CDo TABLE 4-5: DRAG BREAKDOWN: JENSEN METHOD From Jensen's Method the percent contribution of parasite drag for each component was determined.

Interference(6% Landing Gear(44%) Vertical Tail(4%) 4- 14 4.3.3 Landing Gear Buildup Method As illustrated in Figure 4-2, the landing gear is the largest contributor to the parasite drag. Therefore in order to get a better estimate of the landing gear drag, a simple scheme was devised. The landing gear consists of three struts and three tires.

The tires were treated as a combination of spheres (Cd = .4) and cylinders (Cd = 1.0) which gave an average Cd = .7. By modeling the struts as bluff body cylinders of known Cd = 1.0, a better estimate of the landing gear drag could be found as shown in Table 4-6.

Component Cdo A_ (sq.i_) CDo Main Struts 1 0.0063 0.001 Nose 1 0.0025 0.0004 Tires(3) 0.7 0.0558 0.0062 Landing Gear Total 0.0076 0.0127 Rest of Plane (From Nelson Method) Interference 15% Total CDo 0.021 TABLE 4-6: DRAG BREAKDOWN: LANDING GEAR BUILDUP This final method seems to give the most appropriate Coo, since it combines both low Reynolds number effects and a detailed landing gear drag buildup.

4.3.4 Induced Drag In order to find the drag polar for the aircraft, only the aspect ratio and Oswald efficiency factor remain to be calculated. The aspect ratio was fixed at 7.2 for reasons previously discussed. The Oswald efficiency factor can be estimated using component buildup techniques similar to those used for the parasite drag.

1 1 1 1 es.¢ ewiug efus e_r Once again, using the methodology given by Jensen, c,. c can be found. The value for e_,gcan be found from Jensen(Figure 3.3) and is equal to .88. This leaves e_og =.88. The 4- 15

value for the fuselage efficiency canbe found using c_, = • Efus is equal to .6 for

aircraft if Jensen (Figure 3.4) is applied. Assuming eother = 20, the total aircraft efficiency factor, eac = .82 for The Bullet.

4.3.5 Drag Polar The drag bucket for The Bullet is presented in Figure 4-4.

0.18 I I 0.16 ' i 0.14 , _ I = 0.12 o ! _E 0.1 ! O 0.08

a 0.06 . _,._,_ "ii .c0

0.04 ' -e-- Cd (flaps @ 20 deg) 0.02 " "J'_ .. ; - .

& I

0 I .... I,,I ...... i,.,J .........

0 0.21 0.4 0.6 0.8 1 1.2 1.4 1.6 Lift Coefficient

/

CI @ Cruise FIGURE 4-4: AIRCRAFT DRAG POLAR 4- 16

4.3.6 Aircraft L/D

Curve

Max L/D = 14.9 A a "-1 10- m. L/E) _ cruse : 8.5 O .i

iI

rr

/

_ 4-

° t

(

-J R -2. i -4. _ -8 -6 -4 -2 0 2 4 6 8 10 12 Angle of Attack (deg) FIGURE 4-4: AIRCRAFr LIFT TO DRAG CURVE 4.3.7 Drag Reduction Possibilities The drag on the aircraft is a key player in determining whether or not the design cruise speed objective of 55 feet/second could be met. Therefore, drag reduction possibilities were examined with the hope of attaining faster speeds. Three specific techniques were considered: Splitter plates on struts Wheel pants or cowlings Winglets Because the landing gear accounts for about 50% of the aircraft drag, it was targeted as a possibility for drag reduction. The size of the struts were constrained by structural considerations, and tire size was not flexible since The Bullet needed to be equipped for rough field takeoff roll. However, if the struts could be made more aerodynamic in shape by perhaps adding a piece of balsa to act as a splitter plate, the drag could be decreased. This idea may be used in the design of The Bullet when flying indoors 4- 17 rather than outdoors where the plates may fall off in high grass. The idea of changing tires for the indoor and outdoor course was also a possibility, but due to the increased cost of buying an extra set of tires this idea was not deemed feasible. Another option was to use wheel cowlings for the tires. This idea was eliminated solely due to the fact that wheel cowlings are not sold locally, and would most likely hinder takeoff when used on the grass runway. Given more time and analysis, wheel pants technology, or perhaps even a splitter plate technology, could possibly be incorporated into future designs of The Bullet.

The Reynolds number at cruise is approximately 350,000 for the aircraft. Because of this relatively low Reynolds number regime the drag is mostly parasite drag rather than induced drag. Therefore, winglets, which reduce the induced drag, would simply add a weight penalty without much drag reduction. Once again, time constraints eliminated an in-depth study of winglet effectiveness.

4.4 Summary of Aerodynamics The aerodynamic design of The Bullet, especially the wing design, was the most difficult task in the design process. In summary, the high-speed versus short takeoff distance was the primary driving factor in the design. The final aerodynamic parameters are summarized in Table 4-7. Although we are confident in our design, there are some potential problems. First, the flaps may not be fully effective due to the large drag increase. Second, the monokote may deform the airfoil shape and degrade performance. Third, many of our predictions are based on educated guesses from analytical data. For example, the lift curve slope of the airfoil is an estimate from the graph, yet it greatly affects all other aspects of the design. Finally, if our weight exceeds 4.6 lb., then the we run the risk of exceeding the takeoff distance objective of 25 feet.

4- 18 Aircraft Airfoil Flaps 20% c size FX 137-63 Max CL 1.5 50% b Section C1 max 1.6 length CL alpha .077/deg CDo .021 plain type C1 alpha .1/deg e .82 Cmo -.24 L/D max 14.9 Table 4-7 Aerodynamic Summary 4- 19

5.0 PROPULSION

5.1 General Overview Electrical propulsion systems are considered state-of-the-art in Aeroworld, and will be used to power the RPV. The propulsion system is composed of three main components - engine, propeller, and fuel system (batteries). All components function together, and must be selected as a unit. The driving factors behind the propulsion selection process were maximum obtainable level velocity and satisfactory takeoff performance, while incurring the smallest cost and weight penalties. While these drivers were paramount, the other performance design requirements and objectives could not be neglected.

5.2 Propeller Design A code, PROP123, was used to predict propeller performance. The code calculates performance using simple blade element theory. Induced velocity and tip loss corrections were available through the program, and both were employed. Reynolds number and Mach corrections were also options for correcting the airfoil section C! and Cd data, but these refinements were not used.

With only Reynolds and Mach corrections the propeller data was found to be the most conservative, and in that sense the best, estimate of propeller performance.

One restriction was placed on the propeller by the design team. The propeller diameter was not to exceed 12 inches. This was imposed to limit the length of the landing gear, and meet the clearance objective for rough field takeoff without incurring substantial drag penalties. Landing gear, especially thick struts, adds significantly to the overall drag, and diminishes the performance of the aircraft.

Propellers ranging in diameter from 9 to 12 inches were studied. These propellers had 2 or 3 blades and pitch values ranging from 4 to 8 inches, 5-1

depending on availability. From initial studies of the data it was seenthat

takeoff and high speedperformance improves with increasing propeller

A chart outlining this trend is presented

diameter, pitch, and number of blades.

as Table 5-1 below.

Disadvantages

Advantages

length of landing gear

takeoff and high speed

Larger propeller

increases

diameter

improvements

none

takeoff and high speed

Larger propeller pitch

improvements

cost and weight increase

takeoff and high speed

More propeller blades

improvements

Table 5-1: Propeller Selection

An important note from Table 5-1 is that 2-bladed wooden propellers are preferred because they are significantly less expensive ($3.50 as compared to $10.00), and weigh but a fraction of their 3-bladed plastic counterparts (approximately 0.6 ounces less).

The performance of possible propellers was examined, and appears in three graphs. All of the propellers in these charts are 2-bladed except for the 10-8 model which has 3 blades. Fignre 5-1 shows the dependence of propeller efficiency on advance ratio. Figures 5-2 and 5-3 depict the effect of advance ratio on CT and Cp for various possible propellers.

5-2 0.900 -.

Zinger 12-8 I_ Ih,.d n >,0.800 - o 0.700 -':. 10-8 3-blade c- mJ_ "qp •_- 0.600-!

m,- - J J "-- !

"- 0 500 J J w,..0.400-!

/ 0.300-': _-0.200-!

&'- 0.100-!

0.000 J O O O O O O O O O O O O O O O O O O O O O _-- 04 03 _ LO t,D I_ CO Ob o o o o c5 o o 6 o o Advance Ratio, J Figure 5-1: Effect of Advance Ratio on Propeller Efficiency 0.160 - -I-- Zinger 10-7

%

0.140 Zinger 11-7 0.120 Zinger 12-8 0.100 C5 0.080 . 10-8 3-blade 0.060

0.040

0.020 '%"

0.000 O O O O O O O O O O O O O O O O O4 03 _ LO £0 I_ CO Ob o o d o o d o o Advance Ratio, J Figure 5-2: Effect of Advance Ratio on CT 5-3 0.090 - • Zinger 10-7 0.080 • Zinger 11-7 0.070 • Zinger 12-8 0.060 • 10-8 3-blade a.O.050 o 0.040 0.030 0.020-

Y/

0.010-

t \

0.000- O O O O O O O O O O 0 0 0 0 0 0 0 0 0 0 0 _ Cx.I 03 _" _ f_O I"-- CO Ob ddoododd oo Advance Ratio, J Figure 5-3: Effect of Advance Ratio on Cp After considerable study, a Zinger 2-bladed wooden propeller was chosen.

The propeller is the 12-8 - having a diameter of 12 inches, and a pitch of 8 inches.

At first, this might appear to be ill-advised decision. The 3-bladed 10-8 propeller has comparable efficiency to the Zinger 12-8, and larger CT values for a given advance ratio. It might seem that the 3-bladed 10-8 would be the propeller of choice. However, one must recall that thrust, not CT, is the measure of merit because it is directly linked to takeoff performance. The equation for thrust is 2 4 thrust = CTPn dprop • As seen in this equation, the diameter of the propeller greatly affects the resulting thrust at a given RPM and CT value. In fact, thrust, with all else being equal, is 52% less for the 10 inch propeller than the 12 inch propeller.

The power requirements to turn the propeller must also be studied. Cp values for the 3-bladed 10-8 propeller are approximately twice those of the Zinger 12-8 at a given advance ratio. However, power is defined as 5-4 Power = Cppn3dprop 5.

This equation shows that the power consumed by the propeller is a strong function of RPM and diameter. RPM, or n, in the above equation is assumed to be constant in the analysis. This is not exactly correct (assumes torque of each propeller is the same), but is a good approximation. A small change in propeller diameter, however, can greatly affect the power requirement. As diameter increases, so too does the necessary power. This seems to suggest that a smaller diameter propeller would be desirable from the standpoint of power. One should be careful before arriving at such a rash conclusion.

The power produced by the propeller (power available) is equal to the product of the propeller efficiency and power output of the motor. It is desired to have the lowest power output from the motor, but still enough power so that high speed and climb performance can be improved. The Zinger 12-8 propeller produces enough power to satisfy the high speed requirement, while the 3- bladed 10-8 propeller does not. Therefore, from the standpoint of power, as with thrust, the Zinger 12-8 is a better selection than the 3-bladed 10-8 model.

Graphs of thrust and power, rather than their coefficients, would have been more informative. They were not presented because there was a problem with the PROP123 code. The code was repaired, but time did not permit the reproduction of all graphs. Preliminary studies did reveal increased thrust and power required with the Zinger 12-8 propeller over all others, particularly the 3- bladed 10-8. The preliminary performance estimates, while inaccurate, are believed to accurately predict trends. On the basis of the thrust predictions and propeller power requirements (and resulting power available), it is thought, and 5-5 data supports, that the Zinger 12-8 propeller will outperform the 3-bladed 10-8 propeller in the critical areas of takeoff, high speed, and climb.

5.3 Motor Selection The motors considered for use in the RPV were all Astro motors. Astro models 05, 05 FAI, 15, 25 were stocked, and their use was recommended by upper level management. The Astro 05, and 05 FAI were not studied extensively because their low power rating would not allow the high speed objective to be reached. Therefore, only the Astro 15 and 25 models were researched. A comparison follows.

Astro 15 Astro 25 Motor Weight (ounces) 7.5 11 Motor Cost (dollars) 107.00 174.00 Cost/Weight Ratio $14.27/ounce $15.82/ounce Table 5-2: Motor Comparison Preliminary study proved that the Astro 15 outperforms the Astro 25 in the areas of takeoff and high speed for the size of propeller used in the design Only with propellers which produce a large torque load (large diameter, large pitch, and increased number of blades) will the larger Astro 25 motor produce a higher maximum velocity than the Astro 15. Besides better high speed performance, the Astro 15 is also less expensive, and weighs much less than the Astro 25 (See Table 5-2). The Astro 15 is also capable or providing sufficient takeoff performance, and the range requirement can be easily met. These performance features will be explicitly delineated in section 8.0. For these reasons, the Astro 15 was clearly the engine of choice. Two models of the Astro 15 were readily available: one with a gear ratio of 31:14, and the other with a gear ratio of 31:13.

5-6

The model equipped with a gear ratio of 31:14was chosenbecauseit delivers a

slightly higher propeller RPM, and maximum speed. SeeTable 5-3 for the

specifications of the Astro 15motor.

Name Astro 15

Maximum Power 200Watts

Internal Resistance 0.12Ohms

Kv 1.098inch-ounce/amp

Kt 7.94E-4Volts / RPM

Tloss 1.37 inch-ounce Gear Ratio 31:14 Table 5-3: Motor Specifications 5.4 Engine Control and Fuel The fuel for the aircraft consists of 13 nickel-cadmium batteries with a capacity of 1000 mah. Each battery has a nominal voltage of 1.2 Volts, bringing the total voltage to 15.6 Volts (this is maximum allowable voltage for the Astro 15). The high voltage of the batteries provides for a higher maximum velocity, improved climb capabilities, and better takeoff performance. The effect of number of batteries on the maximum velocity and takeoff performance is shown in Figure 5-4. The effect of number of batteries on maximum rate of climb is plotted in Figure 5-5.

5-7 A 9O 3O 'ID Vmax (ft/s) t- O Takeoff O O 28 Distance (fl) O O O v O 80" v O O ¢- 0m 4.J > 70" O E O E i-- 60 , 20 9 10 11 12 13 8 14 Number of Batteries Figure 5-4: Effect of Number of Batteries on Maximum Velocity and Takeoff Distance A _30 n R/C (ft/s) O 2_ E o10' E E "_ 0 I ' I I I I m 9 10 11 12 13 14 _ g Number of Batteries Figure 5-5: Effect of Number of Batteries on Maximum Rate of Climb 5-8 The capacity (1000 mah) only affects the range and endurance of the aircraft. The range requirements for the aircraft were not difficult to achieve (16000 feet), and could have been achieved with batteries of smaller capacitance (600 mah). These batteries would have been smaller, and a possible weight savings could have been garnished. However, the 1000 mah batteries were the smallest available, and will be used in the technology demonstra tot.

The voltage input to the motor will be controlled with a Tekin speed controller. The maximum voltage of 15.6 Volts will be used in the takeoff, climb, and maximum velocity flight configurations. However, the voltage supplied to the motor will have decreased so that level flight can be achieved at speeds between Vstal 1 and Vmax. Only then will the power available and required terms match - the necessary condition for level flight.

At cruise, for example, the aircraft will not require the 15.6 available Volts.

In fact, only 9.82 Volts are necessary. This is approximately 63% of the available throttle. Such information is valuable, and allows the pilot to select the necessary throttle setting on the radio controller to achieve the cruise condition.

5.5 Manufacturing and Installation It is required than the complete propulsion system be removed or installed in less than 20 minutes. This is so that the equipment can be used by several RPVs in the same air show. In order to achieve this goal, the batteries and engine will be made readily accessible. The batteries will be contained with heat-shrink plastic and reside in the wing carry-through structure below an access hatch located on the top of the aircraft. The battery pack will be attached with Velcro, rather than screws to facilitate removal and installation. The Velcro connection will be strong to help minimize any battery translation. Any motion 5-9

could have pronounced effects on the aircraft becausethe batteries represent a

large weight item (over one pound).

The engine is also accessibleat the noseof the aircraft. There will be room

between the fuselage and engine to allow for air circulation, and necessary

cooling. The nosecone sectionwhich houses the engine will also be hinged to

streamline motor installation. An engine mount will be securely fashioned to the

frame of the fuselage, bolted into a sturdy plywood bulkhead. The motor can be

installed or removed from the mount with a turn of a fastener. In summary, the

propulsion systemwill be manufactured so that it can be installed or removed

within 20 minutes. The integrity of all propulsion attachments will not be

sacrificed to achieve this goal.

5- 10

6.0 WEIGHTS AND BALANCE 6.1 Weight Breakdown 6.1.1 Preliminary Estimate When the basic concept for The Bullet was chosen, the preliminary sizing for the plane was made, and an initial weight estimate was calculated. Because the weights of the avionics, the different engines, and the available batteries, were known constants, the actual structure of the plane was the largest unknown. The decision to use the Astro 15 engine, along with the voltage and current requirements, led to the selection of 12 1000 mAh batteries as the power supply. The preliminary estimate for the weight of the structural components was made by looking at the component weights of several previous airplanes.

This data can be found in Appendix B. For example, for each wing, the weight per unit area was calculated, and the average of these values was found. This was then used to calculate an estimate of the weight of The Bullet's wings based on the preliminary value for the surface area. A similar method was used to calculate the weights for the fuselage, the vertical and horizontal tails, and the landing gear. However, because the mission requirements for The Bullet differed from previous years, the weight estimates arrived at using this method were not expected to be very exact.

6.1.2 Secondary Estimate In order to make reasonable estimates for the amount of wing area needed and the structural load requirements, a refined value for the weight of the airplane was needed. Using the initial weight estimate, a preliminary structural design was created. From this, the volumes of the monokote, balsa, and spruce 6-1

needed to build the fuselage, wings, and tail, and volume of the steel needed to

build the landing gear were approximated. Sincethe densities of eachof the

building materials was known, individual component weights were calculated

by multiplying the volume of eachmaterial included in the component by its

density. As can be seenfrom Table 6.1,in most cases,thesevalues turned out

Airplane Initial Weight Revised Weight % of Revised Component (pounds) (Pounds) Total Weight Propulsion 0.704 0.735 16.02 Engine 0.438 0.469 10.21 Gear Box 0.094 0.094 2.04 Engine Mount 0.073 0.073 1.58 Propeller 0.100 0.100 2.18 Batteries 1.095 1.056 23.02 Avionics 0.408 0.433 9.43 Servos (3) 0.113 0.113 2.45 Speed Controller 0.111 0.111 2.41 System Batteries 0.125 0.125 2.72 Receiver 0.059 0.059 1.29 0.563 0.280 6.10 Fuselage 0.781 0.800 17.44 Wing Tail 0.250 0.205 4.46 Horizontal 0.125 0.149 3.25 Vertical 0.125 0.056 1.21 0.375 0.621 13.53 Landing Gear Main Gear 0.250 0.410 8.93 Nose Gear 0.125 0.211 4.60 Glue, etc. **** 0.200 4.35 Error Factor 5% 5% **** Total Unloaded 4.384 4.546 **** Payload 0.035 0.050 1.08 Total Loaded 4.419 4.596 **** Table 6.1: Weight Estimation to be fairly close to the preliminary weight estimates. The only major discrepancies occurred in the weights of the fuselage and the landing gear. The reason for the difference in the estimates for the fuselage weight was the fact that 6-2

the fuselageweight to volume ratio turned out to be considerably lessthan the

original sizing. The difference in the landing gear weights was a result of the

longer struts necessaryto accommodatethe increasedtip clearancerequirement.

Because the design for The Bullet called for a lightweight plane with high ground

clearance, the weight percentages shown in Figure 6.1 for the components differed from previous designs in the obvious areas, the landing gear, the fuselage, and the wing.

Glue, etc.(5%) Passengers(l%) Propulsion(17%) Landing Gear(14%) Avionics(10%) Tail(5%) Wing(18%) Batteries(24%) Fuselage(6%) Figure 6.1: Component Weight Percentages The weight of the components was combined with the weights of the propulsion system, the batteries, and the avionics package to complete the estimate. While all of these weights were given, there were several minor changes between the preliminary and secondary values. In order to provide a more exact estimate, the value for the engine weight was changed to the that found when an available Astro 15 was weighed. The number of batteries in the secondary estimate was boosted to 13 from the original 12. However, while the number of batteries changed, the weight per battery for the 1000 mAh battery 6-3

given by the new catalog was 1.3ounces as opposed to the 1.46 ounce weight

given by the old catalog. Therefore, the total weight of the batteries actually dropped in the secondary estimate. Also, weight was added to the avionics package to account for the wire connections and to the total to account for glue and fasteners.

6.2 Center of Gravity Location Once a reasonable weight estimate was made for each of the components, the location of the center of gravity for The Bullet had to be found. In order to do this, a detailed internal layout of the plane was made, showing the positions of each of the components relative to the nose. Ideally, both the forward (unloaded) and the aft (fully loaded) center of gravity positions should be placed so that the static margin of the airplane falls between 20 and 25 percent of the chord. The static margin was calculated from static margin = XNP Xc_ (6.1) C C where XNp is the position of the neutral point as given in section 7.3 and c is the chord of the wing. In order to achieve the desired static margin with a neutral point location of 0.492c, the center of gravity location had to fall between 0.292c and 0.242c. Therefore, the wing had to be placed so that the airplane center of gravity was slightly aft of the quarter chord location of the aerodynamic center of the wing.

To find the configuration that would provide the desired center of gravity location, the positions of the separate components were varied over limited ranges. The main constraints that had to be considered were the lengths of wire available to connect different components, the fact that the batteries were required to be housed in the wing carry-through, and the lengths available for the servo push rods. Because the wing and batteries essentially must be moved 6-4

together, and the sum of the two makes up approximately 40% of the total

weight of the airplane, this was the combination that was most influential in changing the position of the center of gravity. The final configuration, detailed in Table 6.2 and Figure 6.2, resulted in a center of gravity location of 9.76 inches for the unloaded plane, and 10.03 inches for the fully loaded plane. These values, as was desired, are both slightly aft of the quarter chord position of 9.5 inches.

Airplane Weight (pounds) X Position Component (inches) from Nose Engine, Gear Box, 0.635 2.0 Engine Mount Propeller 0.100 -0.5 Batteries 1.056 10.3 Flap Servo 0.038 13.25 Rudder Servo 0.038 15.5 Elevator Servo 0.038 15.5 Speed Controller 0.111 5.75 System Batteries 0.125 8.0 Receiver 0.059 8.25 Fuselage 0.280 14.5 Wing 0.800 9.5 Horizontal Tail 0.149 35.5 Vertical Tail 0.056 35.5 Main Gear 0.410 12.5 Nose Gear 0.211 4.0 Unloaded 4.546 9.76 Payload 0.050 23.0 Fully Loaded 4.596 10.03 Table 6.2: Center of Gravity Locations 6-5 Propeller Engine, Gear Box, Engine Mount Nose Gear Scale (inohes to inohes) _. Speed Controller System Batteries Receiver .iI II -_ Ving Forward CG Aft CG .7 .................. _-.

,I 4

0.25 I 4 Balteries 0 0 Hain Gear Flap Servo Fuselage Rudder and Elevator Servos Payload el Vertical and Horizontal Tails Figure 6.2: Center of Gravity Locations , , 6-6 7.0 STABILITY AND CONTROL 7.1 Stability and Control Requirements The stability and control requirements of this aircraft were among its most crucial. However, because the desired stability and control characteristics can be achieved with many different configurations, the location and sizing of the stability and control surfaces was engineered last in order to fit within aerodynamic and structural parameters. The aircraft was required to be stable and controllable in the three coordinate directions, namely yaw, pitch, and roll. These definitions led to the specific requirements for The Balsa Bullet: • Pitch, or Longitudinal, stability would be accomplished through the use of a horizontal stabilizer aft of the wing and pitch control through deflection of an elevator on the horizontal stabilizer.

• Yaw, or Lateral, stability would be accomplished through the use of a vertical tail aft of the wing and yaw control through deflection of a rudder on the vertical tail.

• Roll stability would be accomplished through the use of dihedral on the wing and roll control through the combination of deflection of the rudder and the dihedral of the wing.

• The yaw and roll control devices needed to allow the aircraft to perform a 60 ft. radius turn at flight speeds of less than 30 ft/s.

7.2 Pitch Stability The need for pitch stability required that if the aircraft was pitched up, it would correct itself by pitching back down to its previous equilibrium position. The pitch angle, 0, was measured from the horizontal to the 7-1

fuselage reference line of the aircraft and, because there was no wing

incidence, was equal to the angle of attack. Because a pitch-up moment which caused an increase in angle of attack was defined as positive, the slope of the pitching moment coefficient versus angle of attack curve was required to be negative for a stable aircraft. The pitching moment coefficient at zero degrees angle of attack (Cm0) was also needed in order to find the total moment so that the aircraft could be trimmed at any flight condition. The desired magnitude of the slope (Cmo_) was chosen using data from previous designs and from data found in Appendix B of Reference 12. The desired value of Cmc_ at cruise (i.e. fully loaded) was chosen to be -1.0 +_ 0.1. There were three major components of the aircraft which contribute to the Cm0 and Cmc_.

They were the fuselage, the wing, and the horizontal stabilizer.

One of the most important measures of pitch stability was the static margin. The static margin was defined as the difference between the neutral point and the center of gravity location as a fraction of the mean aerodvnamic chord. The neutral point was the point at which Cmo_ was equal to zero meaning the slope of the Cm versus o_ curve was zero. The aircraft was stable for a neutral point aft of the center of gravity. Therefore, a static margin greater than zero was found on a longitudinally stable aircraft. The target value for the static margin of The Balsa Bullet was about 0.25.

7.2.1 Fuselage Contribution The fuselage was a destabilizing component of pitch characteristics meaning that its Cma was positive while its Cm0 contribution was negative.

To find these components, Reference 12 suggested Multhopp's method which breaks the fuselage into discreet sections and applies empirically determined 7-2 The following formulas were employed in this relationships to each section.

method.

1_ 1i x=lf - _2-_'1 K'w 2 if)Ax (7.1) - Z_, f(0%_+ Cmof 36.5Sc ,=o X=] t "_ 1 _ 20E.

= -- 2.,wf _ax (7.2) Cm_' 36.5SU x--0 3o_ Because the fuselage was essentially entirely rectangular, the contributions of all the sections were taken to be uniform. The width of the fuselage was constant at 3.5 inches and the zero lift angle of attack was constant at -6 degrees. These formulas were applied at the aft center of gravity location.

The coefficients found by applying this method were Cm0' = -1.23 X 10-'4 Cm, ' = 3.28x10 -3 per radian 7.2.2 Wing Contribution Assuming that the aerodynamic center of the wing was in front of the center of gravity as was the case in all aircraft studied in the data base, the wing contribution to Cma was positive and to Cm0 was negative. The following formulas found in Reference 12 show that Cm_ and Cm0 depended upon the lift generated by the wing and the moment arm from the lift vector to the center of gravity.

Cmo w =Cm.," + CLo. (x_-_- - _) (7.3) Cm... = CL,/xc--*, C X,_)C (7.4) From inspection of these formulas it was easy to see that there were three factors which influenced the wing contributions to pitch stability. They were the airfoil section, the wing geometry, and the placement of the wing relative to the center of gravity. The chosen airfoil section, the FX63-113, had 7-3

a Cmac= -0.12. The chosen wing geometry altered the lift curve slope of the

wing from that of the airfoil as shown in Section 4. The distance from the

aerodynamic center of the wing to the center of gravity of the aircraft was the

moment arm for the aerodynamic forces. The coefficients were

Cm0" = -9.87 x 10-2

C_,o.= 0.203 per radian

for this aircraft.

7.2.3 Horizontal Stabilizer Contribution

The device which provided the pitch stability and whose size and

placement was driven by the stability and control requirements of the aircraft

was the horizontal stabilizer. It provided a negative contribution to Cmc_and

a positive contribution to Cm0. The effect that the tail had upon these

coefficients depended upon the airfoil section and geometry of the stabilizer

as well as distance from center of gravity of the aircraft to the aerodynamic

center of the stabilizer. The interference due to the wing also affected the

horizontal stabilizer contribution. Reference 12 provided development of the

following formulas for the contribution of the horizontal stabilizer to Cmc_

and Cm0.

Cmo, = rlV.CLo,(¢0+ iv,-i,) (7.5)

Cmo, =--'qV.CLo,(1-- d_ ) (7.6)

where

2CLo (iw - OtLo) e 0 = " (7.7) geAR de _ 2Ct.o. (7.8) dot neAR V, = S,l____ (7.9) Sw7 The values for these coefficients were as follows: 7-4

Cm0 ' = 0.142

Cm_ ' -------1.109 per radian

The two major parameters in the horizontal stabilizer contribution are It and St because they are the most controllable from the design standpoint.

The position of the center of gravity relative to the aerodynamic centers of the wing and horizontal stabilizers has a tremendous effect of the stability of a specific design because it affects the wing and horizontal tail contributions.

Therefore, it is important to have an accurate estimate of the center of gravity location before stability and control analysis is attempted. The sizing of the horizontal stabilizer depends upon the constraints placed upon it by aerodynamic, structural, weight, center of gravity, as well as stability concerns.

The pitching moment coefficients for the entire aircraft were found by summing the coefficients of the individual parts.

The initial location of the aerodynamic center of the horizontal stabilizer was at thirty-six inches. However, the span of the stabilizer would have had to be over four feet in order to meet the stability criteria. This was deemed unacceptable due to structural concerns. When the location was moved to forty inches, the span decreased to approximately 2.74 feet which was acceptable. Figure A-6 shows the pitching moment coefficient versus angle of attack curve for the aircraft at the forward and aft center of gravity locations and indicates the angles of attack required in order to have no pitching moment on the aircraft in the absence of an elevator deflection. The required cruise speeds would then be forty-three feet/second for the forward center of gravity location and thirty-nine feet/second for the aft center of gravity location. This curve was crucial in determining pitch behavior of the aircraft at the extreme conditions.

7-5 7.3 Pitch Control 7.3.1 Sizing and Actuation Pitch control for the aircraft was delivered through the use of an elevator located on the horizontal stabilizer. When the elevator was deflected, the lift on the horizontal stabilizer was altered, changing the pitching moment coefficient for the aircraft. Pitch was used in climbing and diving two obviously important maneuvers. Reference 12 developed equations for determining the effect of the elevator deflection on the moment coefficient. This was accomplished through formulas for determining the slope of the change in pitching moment coefficient versus elevator deflection (Se).

Cm_' = -V,rI_CLo ' (7.10) This coefficient was negative, and the range of its magnitude was determined from examining Appendix B of Reference 12 as 0.9___0.2.

In order to facilitate ease of manufacturing, it was determined that the elevator would run the whole length of the span of the horizontal stabilizer.

The percentage of the stabilizer chord which was elevator and was determined such that the control coefficient was within the acceptable range and such that the elevator wasn't too small to raise serious structural and manufacturing concerns. The maximum elevator deflection angle was determined such that the aircraft could be trimmed at any angle of attack in the normal operating flight regime. The elevator would be controlled by a single flexible control rod extending from the servo in the wing carry- through, traveling through the inside of the fuselage, exiting the through the rear of the fuselage, and attaching to the underside of the elevator. Figures 7- 1 and 7-2 show the pitch moment coefficient versus angle of attack at 7-6

multiple elevator deflection angles for the forward and aft center of gravity

locations, respectively.

0.4

-15 degree deflection

0.3

-10 degree deflection

0.2

"o -5 degree deflection

° I 0 degree

deflection 5 degree

5 -0.1 --r]--

deflection

i-0.2

10 degree --0-- deflection

-0.3

15 degree deflection

-0.4

-6 -4 -2 0 2 4 6 8 10 Angle of Attack (degrees) Figure 7-1: Effect of Angle of Attack and Elevator Deflection on Pitch Moment Coefficient at the Forward Center of Gravity Position 7-7 -15 degree deflection -10 degree deflection -5 degree deflection 0 degree deflection 5 degree -El-- deflection 10 degree -O-- deflection 15 degree --A---- deflection

-6 -4 -2 0 2 4 6 8 10 1

Angle of Attack (degrees)

Figure 7-2: Effect of Angle of Attack and Elevator Deflection on Pitch Moment Coefficient at the Aft Center of Gravity Position 7.3.2 Trimming the Aircraft Perhaps the most important use of the elevator was trimming the aircraft. To trim the aircraft means to actuate control surfaces such that the pitching moment is zero at the desired flight conditions. Trimming the aircraft was necessary to maintain these conditions. Unlike yaw and roll, it was necessary to maintain a cruise pitch angle which in the absence of a rudder deflection would result in a non-zero pitching moment. This was due to the fact that angle of attack played a crucial role in the magnitude of the lift on the aircraft. Reference 12 developed equations for the determination of the appropriate elevator deflection angle to trim the aircraft. Another 7-8

coefficient was defined to aid in this process. It was the slope of the change in

stabilizer lift coefficient versus elevator deflection angle.

= S__, I]-CCLo (7.11)

CL_= 5_ This allows for a tidy form for the elevator deflection to trim (Strim).

_uim = Cm°CL_ ff'CmaCL_m (7.12) Cms, CL_ -- Cm_ ,eLse where CLtrim is the aircraft lift coefficient at which trim was desired.

To trim the aircraft at cruise, the elevator was required to overcome the pitching moment at an angle of attack of -3.3 degrees because it was at this angle of attack that lift was equal to the weight for the designed cruise speed.

Because the design center of gravity location was the aft location, the cruise elevator deflection angle was determined using that location. The angle of incidence of the horizontal tail was set such that the elevator deflection angle to trim at cruise was as close to zero degrees as possible in order to reduce drag. The aircraft can be trimmed at any angle of attack, but the cruise conditions were used in determining the appropriate angles because cruise conditions are encountered the majority of the flight.

7.3.3 Rotation at Take-Off Takeoff rotation was crucial in determining the incidence angle of the tail, the angle the fuselage makes with the ground, and angle of attack. Figure 7-3 shows the forces acting on the aircraft at takeoff.

7-9

Lift

c.g 1 a.(

i t/Weight

_ Lif!tail

)

Landing gear Landing gear Weight Rotation Take off roll Figure 7-3: Aerodynamic Forces During Rotation From the diagram, one notices that the lift on the tail will create a pitch up moment on the aircraft during roll since it is mounted at a negative incidence angle. As the angle of attack changes relative to the ground, the angle of attack of the tail plane changes. When the angle of attack becomes great enough that the lift on the tail changes direction to oppose the increased moment due to the increased lift, rotation will occur. The takeoff rotation analysis sets the incidence angle of the tail as well as the position of the landing gear (See Section 9.4). One other note is that the center of gravity location will move back slightly from its original position from the nose at rotation.

7- 10 St 1.5 square feet 5.0 ARt Airfoil Section NACA 0009 Rectangular Planform Shape Se 0.3 square feet 5e (max) _+15 degrees 8e (trim at cruise) 0.05 degrees Table 7-1 : Horizontal Stabilizer Data Cm. 4.307x10 -2 per radian Cm_ -0.903 per radian Gin6 c -0.721 per radian C L_¢ 0.319 per radian Neutral Point 49.2% MAC 22.3 % MAC Static Margin Table 7-2 : Pitch Coefficients 7.4 Yaw Stability An aircraft is determined to be laterally stable if, when perturbed sideways, the aircraft resumes it original position. The yaw angle, [3, was measured from the body x-axis to the velocity vector of the aircraft in the x-y plane and was positive in the clockwise direction. A positive yaw moment is one which causes the aircraft to rotate clockwise. This means that an aircraft is stable if the slope of the yaw coefficient (Cn) versus yaw angle curve is positive. The range for the magnitude of the slope was determined by 7-11

looking at previous designs as well as the aircraft data found in Appendix B of

Reference12. Two aircraft components contributed to this stabilitv coefficient:

the fuselage, which was destabilizing, and the vertical tail, which was

stabilizing.

7.4.1 Wing and Fuselage Contribution

The wing and fuselage provided a destabilizing component to the vaw

coefficient. The magnitude of the contribution was found through a

combination of the geometry of the aircraft and empirical interference and

correction factors. The method used was that laid out in Reference 12. The

following was the formula for the fuselage contribution to the yaw coefficient

versus yaw angle slope.

Sfsif (per degree) (7.13)

C,_., = -k.kR, $7 DC

For this aircraft design the coefficient was:

C,_., =-4.33x10 -4 (per degree)

7.4.2 Vertical Tail Contribution

The aircraft device used to provide yaw stability was the vertical tail.

Its size and distance from the center of gravity were primarily based upon the

stability requirements of the aircraft. The wing interfered with the flow near

the vertical tail, and the magnitude of this interference depended upon the

wing size and geometry. Reference 12 developed the formulas for the vertical

tail contribution to slope of the yaw coefficient versus yaw angle curve.

1 + d_J_

(7.14)

c.,. =

where

Vv_ S,l, (7.15) Swbw 7- 12 1+ 0 724 + 3 061 + sv/s. cos A,4. +04 + 0 oo ARw d ,716, The latter factor accounted for the difference in flow velocity between the wing and the vertical tail as well as the side swish due to vortices from the wings. For The Balsa Bullet, the vertical tail contribution was: C°_, = 0.0760 Again the total coefficient for the aircraft was found from summing the individual components.

The location of the center of gravity of the vertical tail was taken to be forty inches like the horizontal stabilizer. The size of the vertical tail was determined such that the stability coefficient was within the acceptable range and was acceptable to the structures group. The desired range was again found by analyzing values from past designs as well as from data taken from Appendix B of Reference 12. The range of acceptable values was determined to be 0.05+_0.01.

7.5 Yaw Control Yaw control was accomplished through deflecting a rudder on the vertical tail. The rudder was used to overcome forces in various conditions such as cross wind landings, asymmetric power situations, and spins. The rudder was also used in conjunction with wing dihedral to produce roll control especially during turning maneuvers. Rolling during turning provided a more efficient and more comfortable turn. The measure of yaw control was the slope of the change in yaw moments versus rudder deflection angle curve. The method for obtaining this slope was developed in Reference 12. The calculations result in the following relation.

C,_, = -rlVv'_CL,. (7.17) 7-13

Again for manufacturing ease,the rudder it was determined that the

rudder would run the entire height of the vertical tail. The percentage of the

chord which was rudder was determined such that the control coefficient was

within the predetermined acceptable range and met structural requirements.

As with the previous stability and control coefficients, the range of Cnsr was

determined from previous designs and Appendix B of Reference 12 as

0.06_+0.02. The rudder would be controlled by a single control flexible rod

extending from the servo in the wing carry-through, traveling through the

fuselage, exiting the side of the fuselage, and attaching to the side of the

rudder. Figure 7-4 shows the effect of rudder deflection on the yaw moment

coefficient.

Sv 9.5 square feet ARv 3.0 Airfoil Section NACA 0009 Planform Shape Rectangular Sr 0.25 square feet _r (max) +_30 degrees Table 7-3 : Vertical Tail Data Cn_ 5.12x10 "2 per radian Cnb r -4.61x10 -2 per radian Table 7-4 : Yaw Coefficients 7- 14 0.025 -0.025 -30 -20 -10 0 10 20 30 Rudder Deflection Angle (degrees) Figure 7-4: Effect of Rudder Deflection on Yaw Moment Coefficient 7.6 Roll Stability Unlike pitch and yaw stability, roll stability was not attributed to a surface separate from the lifting surface. Instead, roll stability was a function of the wing placement on the fuselage and dihedral. The fuselage of the aircraft could either be stabilizing, if the wings are mounted on the top of the fuselage, or destabilizing, if the wings are mounted on the bottom of the fuselage. This was due to how the lift on each semi-span on the wing is changed during roll. Figure 2.33 of Reference 12 shows the effect the fuselage has on roll stability.

In Reference 12 the measure of roll stability was defined as the slope of the change in roll moment versus side slip angle curve. If this slope (Cll3) was negative then the aircraft is stable. This is because a positive roll moment induces a positive side slip, so a negative roll moment is needed to return the 7- 15 aircraft to its equilibrium condition. Reference 11 provided a formula for this slope.

F b/2 (7.18) CI_ =-2 S--_ !CL. cydy where F was the wing dihedral in radians. The amount of wing dihedral was determined such that the angle would be small enough to have a negligible effect on lift and large enough to provide a roll stability coefficient within the acceptable range. The target range of values for this coefficient was found from data in Appendix B of Reference 12. The chosen range was -0.08+0.02.

7.7 Roll Control Roll control was necessary in order to bank the aircraft during a turn. It could be provided either by ailerons or by a combination of rudder deflection and wind dihedral. The latter method was chosen for use in The Balsa Bullet because of the limitation of three servos and the choice to use high-lift devices. Reference 12 provides an equation which quantified the roll control of an aircraft as the slope of the change in roll moment versus rudder deflection curve. The formula is provided below.

- S_ Z_zc (7.19) Cl_ ' - L_ v S_b_ The magnitude of this coefficient was not as important as its sign. This was because roll control would not need to produce very much power. Any value under between 0.01 and 0.1 was acceptable. Figure 7-5 shows the effect of rudder deflection on the roll moment coefficient.

7- 16 0.008 0.006 u 0.004 o 0.002 o -O.O02 = -0.004- © -0.006- -0.008 -30 -20 -10 0 10 20 30 Rudder Deflection Angle (degrees) Effect of Rudder Deflection on Roll Moment Coefficient Figure 7-5: C,_ -9.49x10 -2 per radian C_, 1.33x10 -2 per radian Table 7-5 : Roll Coefficients 7-17 8.0 PERFORMANCE 8.1 Takeoff Takeoff performance was a difficult parameter to accurately predict because of the uncertainty in the propeller data. Takeoff was predicted using propeller data obtained from the PROP123 program coupled with a program called TAKEOFF. The TAKEOFF program uses thrust values derived from actual motor performance calculations. Ground roll was also calculated using corrected simple blade element theory predictions (Reference 11-10) with a personally developed FORTRAN code. The code was aptly named GROUND ROLL. In GROUND ROLL, the engine thrust was estimated using the same simple blade element theory with knowledge of the maximum power produced by the engine. Both takeoff codes utilized the equations of motion, and a numerical integration sequence. It was hoped that the calculated ground roll distances of the two methods would be comparable, thus increasing the confidence in the results. Table 8.1 shows the comparison (calculated with la = 0.15), and lends confidence that the aircraft will be able to meet the takeoff distance requirement of a 25 foot roll from a prepared field.

TAKEOFF Predictions GROUND ROLL Predictions 2.94 3.12 Static Thrust (pounds) 24.5 25.3 Takeoff Velocity (feet/second) 1.7 1.4 Time to Takeoff (seconds) 21.3 16.5 Roll Distance (feet) Table 8-1: Takeoff Performance 8-1

It is important to note that the takeoff performance of the TAKEOFF

program is believed to be somewhat optimistic. This is because the propeller performance predicted by PROP123 is better than that produced in an experiment conducted at the University of Notre Dame which used an identical propulsion system. Experimental data revealed that the static thrust would equal 2.94, not 3.12 pounds. Despite the discrepancy, there is reasonable certainty that the plane will fulfill its design objective, and liftoff in under 25 feet.

The rough field takeoff requirement was not as restrictive as the indoor requirement. When the value of ]a is doubled (to represent long grass), the ground roll distance is increased by only 5 feet. The takeoff distance from long grass was found to be 26.3 feet, much less than the requirement of a 60 foot maximum ground roll.

8.2 Cruise After climbing to 25 feet, the aircraft will cruise at a velocity of 55 feet/second. This velocity is higher than planes in the current Aeroworld fleet, and provides customers with a high speed alternative. The necessary CL for the cruise configuration is calculated from the equation: Lift = Weight = 2PVcrui_2SCL, and is found to be 0.202. This CL can be obtained at an angle of attack of -3.3 degrees. This is a small angle, and allows for a relatively level attitude in the cruise configuration. This attitude will provide passengers and pilot a comfortable ride. Also, the cruise condition can be obtained with a voltage setting of 9.82 Volts. This is only 63% of the 15.6 available Volts, and gives 8-2 the aircraft sufficient excess power to climb or maneuver. The current draw in the cruise configuration is 7.05 amps.

In the cruise condition, the L/D of the plane is 8.5. The maximum L/D for the aircraft is 14.98 and occurs at a velocity of 31.5 feet/second. From a purely aerodynamic standpoint, the airplane would achieve lower fuel costs if the maximum L/D condition existed at cruise. However, by flying at 55 feet/second the airplane achieves attractive improvements in speed, and economic improvements (in area of depreciation) as well (See 10.0: Economic Analysis). Therefore, a cruise speed of 55 feet/second can be justified.

Higher cruise speeds could also be obtained. At higher cruise speeds, however, the range capability would be lessened, and higher capacity batteries would be necessary. In The Balsa Bullet demonstrator, the range requirement was grossly exceeded because appropriate batteries were not available at time of construction. It is hoped that eventually the design batteries (600 mah) will become available. If the proper batteries are installed, the design cruise velocity of 55 feet/second will be appropriate. As it now stands, higher cruise velocities would not cause exception to the design requirements and objectives, and are recommended.

8.3 Turns The aircraft, a low wing design, will use rudder deflection coupled with a 5 degree wing dihedral to achieve the banking necessary to negotiate a turn.

In order to meet the requirement of a 60 foot radius turn at 30 feet/second, the calculated bank angle must be 25 degrees. This angle was found with the aid of the following relation: tan_p- V_2 gR 8-3 where qbis the bank angle, and R is the turn radius.

In order to evaluate turn performance, it was necessary to study the roll

capability of the aircraft. No formula was found which estimates the roll rate

as a function of rudder deflection and dihedral. To crudely approximate the

steady state roll rate, Pss, an equation involving ailerons was modified and

appears below (References11-7,11-12):

The calculated roll control of 0.0133/rad (C10r), and a maximum rudder roll

deflection of 30 degrees, the aircraft can roll to the necessary bank angle in 5.15 seconds. This is a very long roll time; it is much greater than similar Aeroworld designs of the past. On this basis, it is believed that the roll capability of The Balsa Bullet will be markedly better than that predicted by the crude formula above.

The minimum level turn radius of the aircraft was also computed using Using this equation the minimum turn radius was found to be 28 feet. This turn radius is much smaller than the 60 foot maximum imposed for the Vturn of 30 feet/second.

8.4 Landing The landing performance for the aircraft was estimated using techniques similar to those employed for t_akeoff. The roll distance was 8-4

calculated using the same equations with appropriate initial conditions and

the thrust set equal to zero. To validate the landing roll code, results were compared to those predicted by where, B = 1CDPS A = _tW.

This equation was found in reference 11-10. The landing roll distance calculated from the two methods were 52.5 and 52.6 feet respectively with a friction coefficient of 0.15, and Viand = 24.5 feet/second. The close agreement of the two solutions produces confidence in the results.

The plane must be able to stop in a distance of 40 feet to service any Aeroworld airports. Obviously, with calculated roll distances of over 50 feet, the plane does not meet the objective. However, the RPV is not required to meet this objective because it lacks braking or reverse thrusting capabilities.

In the actual production these features will be incorporated, and the landing performance will satisfy the objective. It should be noted that brakes will add to the overall weight of the aircraft, and will adversely affect the range of the aircraft.

8.5 Power Available and Required A graph of power available and required contains a lot of information.

The graph shows not only the maximum obtainable level velocity, but also the voltage necessary to cruise at various speeds in the flight envelope. This is done by realizing that level flight is achieved when the power available and 8-5

required curves intersect. From Figure 8-1, the maximum obtainable speed of

the aircraft is approximately 84.6 feet/second.

Preq Pav (15.6 Volts) Pav (14.0 Volts) Pav (11.0 Volts) 2°°1 i: Pav (9.82 Volts) Pav (8.0 Volts) 20 40 60 80 100 Velocity (feet/second) Figure 8-1: Effect of Velocity on Power Available and Required 8.6 Climbing and Gliding Upon takeoff, the initial rate of climb is 15.8 feet/second. The maximum rate of climb occurs at a velocity of 45 feet/second, and its value is 20 feet/second. Using the lowest value of rate of climb for the ascent profile, the plane reaches its design altitude of 25 feet in a time of 1.6 seconds. Figure 8-2 shows how rate of climb varies with velocity at the maximum voltage setting of 15.6 Volts.

The plane, starting from rest, must also be able to clear a 50 foot obstacle in under 200 feet. With ground roll and climb, The Balsa Bullet will obtain a height of 50 feet after only 89.4 feet have been covered. Therefore, the plane easily meets the climb requirement.

8-6 2O [] P,/C (fvs) A o E o 0 j i • 40 60 80 Velocity (feet/second) Figure 8-2: Effect of Velocity on Rate of Climb Another graph was also produced using rate of climb data (Vactual = 15.6 Volts). Figure 8-3 was made because it provides very useful performance information. By graphically showing the effect of horizontal velocity on rate of climb (R/C), the largest possible climb angle and the climb angle which produces the greatest rate of climb can be found. These angles were found using the following relation: Oc = _,H. Velocity,) The maximum angle of climb was found to be 38.4 degrees, and occurs at a horizontal velocity of 20 feet/second. At a horizontal velocity of 40 feet/second the angle of maximum climb is calculated to be 26.6 degrees.

Besides providing data regarding the climb angle, Figure 8-3 also shows the maximum obtainable level velocity. The highest velocity is the horizontal velocity when the rate of climb is zero. This occurs at a velocity of 8-7

approximately 84.6 feet/second, and is in close agreement with velocity

information obtained from power available and required curves.

2O [] R/C (ws) A "1_ "10, E o 0 | I ! ' !

0 20 40 60 80 Horizontal Velocity (feet/second) Figure 8-3: Effect of Horizontal Velocity on Rate of Climb In the event of catastrophic engine failure, it is imperative that the valuable components of the aircraft survive. Excellent glide performance will help ensure this. The minimum glide angle was calculated using the following: =(½)=.

where 7 is the glide angle. With the maximum L/D ratio of 14.98, the minimum glide angle was found to be 3.82 degrees. With this glide slope angle the plane is able to negotiate a horizontal distance of 15 feet for every foot of vertical height. With this glide capability, the airplane can easily avoid any serious damage which might otherwise be incurred to the propulsion system or avionics package.

8-8 8.7 Range and Endurance Simplified equations were used to calculate the endurance and corresponding range. Endurance was found by using 970mah Endurance = ia where 970 mah is the capacity of the battery pack allowing for 30 mah to be drained during taxi. ia represents the current draw for the cruise velocity of 55 feet/second. Range is calculated using the endurance as Range = Endurance * V._,.

The effect of velocity on range and endurance is shown in Figure 8-4.

I0oo Range (ft) 30000 t Endurance (s) A "o 28000 t e- o 26000 t ' 600 o o

t

e-

1 400

ul 20 40 60 80 1O0 Velocity (feet/second) Figure 8-4: Effect of Velocity on Range and Endurance For the cruise configuration the range was found to equal 27241.5 feet, and the plane could remain airborne for 8.26 minutes. The maximum values for 8-9 endurance and range were found by calculating both quantities at every velocity of the plane's flight envelope; from Vstall to Vmax. The results appear in Table 8-1.

Endurance (minutes) Range (feet) Velocity (feet/second) 13.92 25055.1 30 (Maximum Endurance) 45 (Maximum Range) 28692.2 10.63 8.26 27241.5 55 (Cruise Condition) Table 8-1: Range and Endurance Values A more detailed calculation of range and endurance for the actual flight profile was also performed. The results follow in Table 8-2.

Time (minutes) Current Draw (Amps) Flight Regime Range (feet) Takeoff 0.023 16.70 (6.5 mah) 0.026 63.1 18.41 (7.9mah) Climb (25 ft/s) i7.510 24783 7.05 (881.8 mah) Cruise (55 ft/s) 1.000 1800 4.43 (73.8 mah) Loiter (30 ft/s) 0.000 0 0 (0.00mah) Landing (25 ft/s) Totals 8.55 minutes 26646.1 feet 970 mah (30 mah-taxi) Table 8-2: Detailed Range and Endurance for Flight From this table is seen that the maximum endurance for a flight involving takeoff, climb, cruise velocity, and one minute loiter is 8.55 minutes. In the flight, the plane will travel 26646.1 feet. These values are comparable to those calculated under the assumption that the cruise condition was maintained for the entire flight.

8- 10 8.8 Ceiling The absolute and service ceilings for the aircraft were also calculated.

In Aeroworld there are no large mountains to negotiate, and it is assumed to be at sea level. However, it is important to know the altitudes at which the airplane can effectively function. From Figure 8.5, the absolute ceiling is seen to be 52177 feet, and the service ceiling (where R/C = 1.65 feet/second) is at 47872 feet.

A "o [] R/C (ft/s) c o ,1o E I " I I " I I-- " 0 10 20 30 40 50 60 Height (lO00s of feet) Figure 8-5: Effect of Height on Rate of Climb 8- 11 9.0 STRUCTURAL DESIGN The Long Shot Aeronautics structural design philosophy was to provide for the structural needs and requirements of The Balsa Bullet with a factor of safety of 1.5, while attempting to achieve the lowest structural weight in a relatively simple-to-construct design. Cost was also an issue. Because manufacturing labor expenses are a larger percentage of the total cost of the aircraft than raw materials, the emphasis of the design was placed on ease of manufacturing and not on raw material cost.

9.1 Loading 9.1.1 Flight Loading The most strenuous maneuver which The Balsa Bullet was designed to accomplish was a sixty foot radius turn at a speed of thirty feet per second. This maneuver would place the aircraft at a twenty-eight degree bank angle from level flight, and incurs a load factor of 1.13 on the aircraft. Using a factor of safety of 1.5, this would have made the limit maneuvering load factor 1.7, which can be associated with a bank angle of fifty-four degrees. A sixty degree bank angle corresponds to a load factor of 2.0, and this was viewed as a more conservative estimate of the limit maneuvering load factor which The Balsa Bullet might encounter in flight. A negative limit maneuvering load factor of one was adopted, guided in part through the knowledge that The Balsa Bullet was not designed to fly in extreme dive maneuvers, and in part by standard Federal Air Regulations (F.A.R.). The resulting V-n diagram for The Balsa Bullet is shown in figure 9-1. Gust loads were not examined for The Balsa Bullet. Future designs should analyze these, as they may prove significant for outdoor flight.

9-1 2_ 1.5"

/

O <-V st all _3 0.5-

/

-0.5 V max -> 0 10 20 30 40 50 60 70 80 Velocity, feet per second Figure 9-1: V-n Diagram For this flight envelope, the maximum bending moment in the wing was computed to be 165.6 inch-pounds, occurring at the root of the wing.

Additionally, the maximum bending moment in the fuselage was found to be 25.5 inch-pounds occurring at the wing/fuselage joint location.

9.1.2 Ground Loading The free body diagram in figure 9-2 shows the ground forces on The Balsa Bullet in a stationary, fully loaded configuration. The force occurring on the main landing gear is assumed to be distributed evenly on each of the two wheels (1.5 pounds on each wheel).

9-2

_cg

_r

** Note: Wing and

1.6 lbs 4.6 lbs

empennageleft out

for clarity.

3 lbs

Figure 9-2: Aircraft Free Body Diagram -- Ground Loads Based on previous Aeroworld design techniques, a landing load factor of three was adopted for the design of the landing gear for The Balsa Bullet. If the configuration in figure 9-2 is considered to be a landing load factor of one, a landing load factor of three would be three times each force acting on each respective landing gear assembly (that is 4.5 pounds on each of the main gear assemblies). The factor of safety of 1.5 required each main landing gear assembly to be designed to withstand a load of 6.75 pounds. Landing loads were assumed to occur on the two main landing gear assemblies only. (That is, a two point landing was the only scenario considered.)

9.2 Materials Materials selection was based on the strength requirements for primary members and on weight and cost for non-critical members. Spruce, balsa, ca rbon fiber, and steel were selected for use in structural applications. Aluminum was also considered as a candidate, however the availability and more simple 9-3

manufacturing qualities of carbon fiber showed it to be a superior candidate in

high strength situations. Table 9-1 provides a comparison of all materials

considered.

Material

Compression (psi) P (lb/in3) E(psi) x 10^6

c_ Tension (psi) Balsa 400 600 0.009 0.065 Spruce 6200 4000 0.016 1.3 Carbon 130000 126000 0.058 24 Steel* 21000 (shear) 36000 (y) 58000 (u) 0.284 29 Aluminum 20000 15000 0.100 10 *Note: (y)-->yield, (u)-->ultimate Table 9-1: Properties of Selected Materials 9.3 Wing Design The design of the wing for The Balsa Bullet was driven by the high speed objective at cruise and the take-off distance requirement. Compromising the two factors necessitated the use of half-span flaps, and resulted in a relatively high wing loading.

With the half span flaps using 20% of the chord, there was an opportunity to make a three spar wing construction, with spars located at the leading edge, 25% chord, and 80% chord locations. Because the FX 63-137 airfoil section is thin at the 80% chord location, a two spar design was attempted, as the spar at the 25% chord location had a much larger potential cross-sectional moment of inertia, and would likely be able to carry the primary loads by itself. In addition, the two spar design was less complicated in both analysis and construction. (A half-span bulkhead was designed for use at the 80% chord in order to provide for flap attachment.)

For the wing design, the primary load bearing member was the 25%chord

main spar. Figure 9-3presentsthe wing lift, shear,and bending moment

distributions in the spar at a load factor of 2.

0.14-

0.12

-_ 0.1

"/3 \ 0.08 " = _0.06 " "" 004 • ,-,_ .

0.02-

o-

0 5 10 15 20 25 30 35 40

f

-0.5

,_--'I

J "_ -1.5 f f = -2 f J _-2.5 f

"2-3

-3.5 -4 ¢d3

-4.5 -/"

-5 0 5 10 15 20 25 30 35 40 90-

_0-

¢70- N

x

_o- \

_O-

ZL'30-

_0-

_I0- N..._._

_0-

0 5 10 15 20 25 30 35 40 Spanwise Position (inches) Figure 9-3: Half Span Wing Loading for Load Factor n=2 9-5

Figure 9-4 shows the main spar crosssection chosen for use in the wing of

The Balsa Bullet. With a factor of safety of 1.5, the spar was designed to withstand a bending moment of 248.4 inch-pounds at the root. The carbon fiber spar caps were selected for use in stiffening the wing. Figure 9-5 shows partial results of a trade study conducted on the main spar design. Note that at the design point, where the spar cap thickness is 0.25 inches, wing tip deflection was reduced by 80% due to the carbon fiber. The main spar has a margin of safety of 5, due to the significant strength improvement given by the carbon fiber.

_(0.25" x 0.0325") _ _.4_Carbon Fiber Spar Cap "NSpruce Spar Cap 1.52" (0.25" x 0.25") Figure 9-4: Main Spar Cross Section 9-6

0.7

\

0.6

Vithou Carbc n Spar Caps

0.5

U r' ._ v 0.4 o u 0.3 _ 0.2 b- 0.1 WiG Carb( n Spar Caps 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 Spruce Spar Cap Thickness (inches) Figure 9-5: Wing Tip Deflection Analysis 9.4 Landing Gear 9.4.1 Considerations The design philosophy regarding The Balsa Bullet's maneuverable nose wheel tricycle landing gear was to create main gear struts that would yield without failing in an extreme landing situation. Minimization of both drag and weight were considered critical. Landing gear design was constrained by the rough field requirement (three inch grass) in terms of strut length (constrained by tip clearance and propeller diameter), and wheel diameter, as well as the elimination of using of a shock absorbing tensile wire between the main struts. A rough field was also thought to interfere with the ground operation of a tail- dragger configuration. The fixed diameters available for steel strut rods further limited the design choices.

9-7

9.4.2 Strut Selection

Stand-alone struts were selectedfor the main landing gear design, as

opposed to landing gear struts connectedby a shock absorbing spring or wire.

This was due primarily to the three inch clearancerequirement. A shock

absorbing wire was thought to causemaneuvering difficulties in rough grass.

Becausethe density of steel is high, a seemingly small increasein the main

strut diameter resulted in a large weight increase. This caused the design

landing load factor of three to be called into question. A specific concern was

whether a design based on a load factor of three in addition to the required factor

of safety of 1.5 would create landing gear which would be too stiff. The extreme landing case was also questioned on the grounds of probability -- what are the chances that an aircraft would actually impact at a load factor of three?

Although the questions raised were not examined in detail (perhaps they should be), significant landing impacts have been observed in Aeroworld (Diamond Back, 1993). Due to these occurrences and the desire to be cautious in the event of an upset, the decision was made to design the landing gear at the previously decided load factor of three.

A steel strut of diameter 5/32 inch was found to yield without failing under the required loading conditions, providing a margin of safety of -0.24.

This negative margin is justifiable, as the landing gear has been designed to deflect and absorb some of the forces of a hard landing, much like a spring, rather than transmit the force of impact through to the main spar. Figure 9-6 shows a sketch of the main landing gear attachment to the main lower spar cap.

9-8

Rib Section

5/32 "Steel Rod

Foam Tire

! \

Birch Plywood Flooring Lower Spar Cap Figure 9-6: Landing Gear Detail The nose strut was selected to be a 5/32 inch diameter steel rod. The steering assembly and strut will be off-the-shelf and will attach to the main plywood floor of the aircraft, just aft of the motor mount bulkhead. These "pre- fabricated" components were selected with and eye towards the associated time (and hence cost) savings which they should provide. Steering control will be provided by the rudder control servo.

9.4.3 Wheel Selection The primary drivers in wheel selection were weight, drag considerations, and the wheel's ability to roll in three inch grass. While the third factor was 9-9

largely intuitive, it did serve to narrow down the tire choices. A two inch

diameter wheel was believed to be able to roll in three inch grass, however it was not believed to be able to overcome the starting friction in the grass. A 2.5 inch diameter wheel was the smallest diameter tire which was believed to be able to overcome the initial friction of the rough field and was selected for use on all three struts. Foam tires were selected, as they are 15% lighter than rubber tires with similar diameters. Larger diameter tires were desirable for easier take-off and ground handling characteristics, however, a tire diameter reduction of 0.75 inches (from 3.25 inches to 2.5 inches) reduced the total aircraft drag by approximately 5%.

9.5 Fuselage Figure 9-7 shows the fuselage for The Balsa Bullet. The length was limited by the necessary payload volume on one extreme, and by the necessary length for stability and control on the other (while maintaining reasonably sized empennage surfaces). Weight was also a consideration, and spruce was limited to high stress areas.

9- 10

Plywood Engine

Mount Location

3.5"

\

Wing Spar Location Plywood Floor a) Side View m 3.5" Balsa Plywood Floor Spruce Plywood b) Bottom View Figure 9-7: Fuselage The main longerons of the fuselage were designed to withstand a bending moment of 38.25 inch-pounds. The 1/4 by 1/4 inch spruce longerons selected provide a margin of safety of 2.0 in this loading configuration.

9.6 Empennage The empennage design featured the use of NACA 0009 airfoil sections for both vertical and horizontal surfaces. For manufacturing simplicity, the control surfaces were ideally designed to be solid balsa sections with cut out holes for weight savings. Due to availability concerns, though, the rudder design had to utilize spar and rib construction. Figures 9-8 and 9-9 show the vertical 9- 11

tail/rudder and the stabilizer assembliesrespectively. Due to control horn

mounting considerations as well as rib/spar joining considerations, thicker balsa spars were used in place of slimmer spruce spars. Both sections use all balsa designs.

, a) Top View Nylon Hinge 2.45" 2.45" b) Side View Figure 9-8: Vertical Tail Assembly 9- 12 5.28" a) Side View b) Top View Figure 9-9: Stabilizer Assembly 9- 13 10.0 Economic Analysis 10.1 Economic Requirements and Objectives Long Shot Aeronautics began with the goal to create a high-speed, low cost plane to compete in an Aeroworld market that was sagging. In order to compete in this market, the total aircraft cost to the consumer became a primary objective for The Bullet design. The primary goal was to create an affordable aircraft that could be mass produced without sacrificing cruise and takeoff performance. The specific economic requirements and objectives as determined by Long Shot Aeronautics were: - Maximum manufacturing cost of $1600 -Maximum labor cost of $900 (90 hrs of construction) - minimize waste and associate costs - maximum limit of $290.00 on raw materials -minimize cost per Jlight 10.2 Cost Estimates The total estimated cost for The Bullet is laid out in Table 10-1. Based on this estimate, all the economic objectives were met. However, this is contingent upon labor costs totaling no more than $900. As illustrated in Figure 10-1, the largest costs occur in the manufacturing process (55%).

10- 1 Waste(l%) Toolin mlsion System(12%) (14%) Labor(55%) Materials(11%) FIGURE 10-1: COST BREAKDOWN OF THE BALSA BULLET In order to meet the labor cost objective, a very detailed and efficient manufacturing plan must be developed. Otherwise, the cost of the aircraft will increase approximately $16.00 for every extra hour of work. Thus, ease of construction became a catalyst in our drive to keep labor costs down.

10.3 Direct Operating Costs The direct operating costs per flight are composed of the depreciation costs, operation costs, and fuel costs. The breakdown of direct operating costs for The Bullet are shown in Table 10-1. These values were calculated using the procedures set forth in Reference 1. Early on in the design process, a range of 16,000 feet was selectedas a design requirement. This allows The Bullet to service over 50% of the flights to any two airports in Aeroworld with a one minute loiter. The decision to choose a relatively short range was very cost effective. Not only could money be saved by using a lower amphour battery, but also the direct operating costs of the Bullet were reduced. For a design range of 10- 2

16,000ft and cruise velocity of 55 feet/second, the design flight time is only 4.8

minutes. The resulting number of flights possible is 1238,which means the

depreciation costsper flight are only $2.07. Fuel costswere directly proportional

to the current draw, which increaseswith velocity, and inversely proportional to

the lift to drag ratio which decreaseswith velocity. The fuel costsrange from

$1.57to $1.93,depending on the price of an amp-hour. The resulting costsper

flight range from $3.82to $4.18. The breakdown of the cost per flight as a

function of cruising velocity is shown in Figure 10-2.

4,5. A 4. \ Desil :n Ra]t2e = [6000 ft _ Depreciation 3.5- _k - "El-- Fuel 3- _ _ Operations

•=-

._2.5- 0 "I 1 -- " O.5- m O- '_ " "- " "- ....

20 30 40 50 60 70 80 Velocity ft/s FIGURE 10-2: DIRECT OPERATING COSTS AS A FUNCTION OF VELOCITY 10.3 Costing Factors The three primary cost factors that will determine the marketability of the Bullet in the Aeroworld market are -cost per flight (CPF) -cost per lO00 ft (CPKFT) -cost per minute (CPFM) 10- 3 The specific formulas to determine these cost factors can be found in Reference 1.

The CPF for the bullet is $3.93, the CPKFT is $.23, and the CPFM is $.76. The economic objective to be competitive in Aeroworld facilitates the need to minimize these cost factors. In order to achieve this goal, one would think that The Bullet should fly at its most aerodynamically efficient cruise speed to reduce fuel costs.

This occurs are 31 feet/second, where the lift to drag ratio is a minimum.

However, upon further analysis as shown in Figure 10-3, this is not the case.

In fact, the optimum cruising speed in terms of cost per flight is 60 feet/second, which even betters the design cruise speed objective of 55 feet/second. Although the cost per flight minute is higher at this speed, the rationale is that the people of Aeroworld are willing to pay a little extra if they can get to their destination a little quicker.

It should also be noted that the maximum attainable range for The Bullet is 27,000 feet due to the fact that smaller batteries could not be purchased.

However, this allows for flexibility in the costing factors of The Bullet. If the maximum range of 27000 feet is used in calculating the costing factors, the CPF turns out to be $5.93, the CPKFT = $.22, and the CPFM = $.1. Comparing these values with the values found using a design range of 16,000 feet, it is clear that for a little extra cost per flight ($2.00 more), the prospective customer can now make over 80% of the flights in Aeroworld. At the same time, the CPFM, is reduced drastically. This flexibility in the aircraft can satisfy a larger portion of the prospective buyers in the Aeroworld market. The Bullet can either be used for short quick flights, or longer more time consuming flights, without an enormous increase in cost.

10- 4 10.4 Economic Summary Low cost was a main focus in the design of The Bullet. However, as the design process evolved, low cost was not deemed as important as meeting high- speed performance and takeoff requirements. Therefore, several decisions made in the design process may not seem very economical. For example, the decision to use flaps, was not very economical. The use of flaps added fixed costs in terms of an extra servo ($35), and they also pose a major construction risk. The decision to use flaps may require extra labor costs as well. Also, the selection of the FX- 63-137 airfoil may pose some economic problems since it may also be difficult to manufacture. The thinking behind these decisions was that the increased aerodynamic performance will actually save the company in the long run both in terms of cost per flight and fuel costs which makes The Bullet more marketable.

10- 5 Fixed Subsystems Propulsion $107 motor $5O speed control $39 batteries (13 panasonic 1000 maph) $5 propeller(Zinger 12-8) Controls radio receiver $35 radio transmitter $75 $10 avionics battery pack switch harness $5 $105 servos(3) $4 wiring Subtotal $435 Raw Materials balsa $30 $3O spruce $5 plywood $50 carbon strips monokote $30 $10 glue miscellaneous $10 $2 landing gear struts wheels $12 Subtotal $179 Manufacturing labor $900 $100 tooling $20 Waste Disposal 1,634 Company's Cost Overhead x 1.4 2287.6 Profit x 1.12 2562.11 $2,562.11 Selling Price Vcruise = 55 ft/s 1,238 Number of Flights $2.07/flight Range = 16000 ft Depreciation Costs $.177/flight Operation Costs Fuel Costs $1.50/maph $1.57/flight $3.00/maph $1.93/flight $3.82-$4.18 Total Cost Per Flight TABLE 10-1: TOTAL COST BREAKDOWN 10- 6 11.0 References 1. Batill, Stephen. "AE441 Lecture Notes." University of Notre Dame, 1994.

2. Beer, Ferdinand P. and E. Russell Johnston, Jr. Mechanics of Materials.

McGraw-Hill: New York, 1981.

3. Design Proposal for The Airplane. University of Notre Dame, 1993.

4. Design Proposal for The Bunny. University of Notre Dame, 1993.

5. Design Proposal for Diamondback. University of Notre Dame, 1993.

6. Design Proposal for Gold Rush. University of Notre Dame, 1993.

7. Design Proposal for The RTL-46. University of Notre Dame, 1993.

8. Dunn, Patrick F. "AE454 Lecture Notes." University of Notre Dame, 1993.

9. Jensen, Daniel T. "A Drag Prediction Methodology for Low Reynolds Number FLight Vehicles." University of Notre Dame, 1990.

10. McCormick, Barnes W. Aerodynamics, Aeronautics, and Flight Mechanics.

John Wiley and Sons: New York, 1979.

11. Nelson, Robert C. "AE444 Lecture Notes." University of Notre Dame, 1993.

11-1 12.Nelson, Robert C. Flight Stability and Automatic Control. McGraw-Hill: New York, 1989.

13. Selig, Michael S., John F. Donovan, and David Fraser. Airfoils at Low Speeds.

HA Stokely Publishing: Virginia, 1989.

11-2

Appendix A: Deliverable Items

Figure A-l: Effect of Weight on Range = Range (ft) 2700E 2680_ O O 2660E G r= ¢0 n- 2640E 2620E 260O( I I 4.5 5.0 5.5 6.0 Total Weight (pounds) A-1 Figure A-2:FX63-137 Airfoil Lift Curve O ""!"'!"'!"'! ...... !"'!"T"! ..... !'"!'"i'"!"" ...... • 0 0 0 , , .......,,..........,.,.,,,.,...,..,, ...... ,°,.,°o°, .......

i ! i i i i ! i i i i i Clmax _ _ _ _ "_ .......

'i!!i .....____ ......_.i_...

_ _oi_oi...i_.i._ :::: ..... ....

_ --i-°i--i°-_ ....i-.i..i..°i .......io°.i.ooi_.i°.o

.... :...,.-!-.i.°.

• °oO°,°,o°.o ....... i°°°•°°°o ....... I°°°l ............... .°°° ...... o .....

............ • ....... m................... i..., ...............

-.-_-.._.-._..-_...,..._°.._°.o_..._.,.,..._..._..._..._...

..._..._..._..._ ...... _..._..°_..._ ....... _..._..._..._ ....

I --10 0 10 20 A-2 Figure A-3: Aircraft Lift Curve With and Without High Lift Devices 1.1 f w/flaps @ 9 deg 1.4- Stall @ 11 deg 1.2

b'

--t-- CI flaps @ (20deg) -0.2 -8-6-4 -2 0 2 4 6 8 10 12 Angle of Attack (deg) A-3 Figure A-4 Aircraft Drag Polar With and Without High Lift Devices 0.18 •-m-- Cd --@-- Cd (flaps @ 20 deg) I I I il I I 1.2 1.4 1.6 A-4 Figure A-5: Lift to Drag Ratio Dependence Upon Aircraft Angle of Attack UD ) 14. y _INm,_.

"'IN_L,, 12.

_o I' "

._o 8. "L/E @ :ruise = 9.7 rr 6- _

_ /

a 4-

° T

-J m

]

-2 -8 ..6-4 -2 0 2 4 6 8 10 12 Angle of Attack (deg) A-5 Figure A-6: Effect of Angle of Attack on Pitch Moment Coefficient at the Forward and Aft Center of Gravity Locations 0.1- 0.05- .l..a .,,-4

\

_-_ 0- • aft center of gravity ©

\

(.3

E

:_ -0.05.

u ,.l..a -0.1- -0.15 position (27.2% MAC) -0.2 -6 4 -2 0 2 4 6 8 10 12 Angle of Attack (degrees) A-6 Figure A-7: Effect of Velocity on Power Available and Required Preq Pay (15,6 Volts) = Pay (14.0 Volts) Pav (11.0 Volts) m _ i Pav (9.82 Volts) Pav (8.0 Volts) $ O 0.

60 80 4O 2O Velocity (feet/second) A-7 Figure A-8: Effect of Advance Ratio on Propeller Efficiency 0.900 Zinger 12-8

0.800 ,11-I • \

>, 0.700- 7 = 0.600- .__ o ._=,,, 0.500- / --_ • 0.400 / 0.300 (3_ 0.200 0.100 0.000 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 ,-" 04 09 ,_" 1.19 (D I'-. CO (:;) o d d o d o o d d d Advance Ratio, J A-8 Figure A-9: Weight and Balance Diagram Propeller Engine, Gear Box, Engine Mount Nose Gear Scale -¢ (inches to inches) Speed Controller Sgstem Batteries Receiver J _/ing For_card CG Aft CG ='_.,.= _ ¢¢

,I 4

0.25 1 4 Batteries O 0 Main Gear -¢ Flap Servo -¢ Fuselage =¢ Rudder and Elevator Servos Pag load Vertioal and Horizontal Tails A-9 Table A-10: Weight Estimation % of Revised Airplane Initial Weight Revised Weight Component (pounds) (pounds) Total Weisht 0.735 16.02 0.704 Propulsion 0.438 0.469 10.21 Engine Gear Box 0.094 0.094 2.04 Engine Mount 0.073 0.073 1.58 Propeller 0.100 0.100 2.18 Batteries 1.095 1.056 23.02 Avionics 0.408 0.433 9.43 Servos (3) 0.113 0.113 2.45 Speed Controller 0.111 0.111 2.41 System Batteries 0.125 0.125 2.72 Re ceiver 0.059 0.059 1.29 0.563 0.280 6.10 Fuselage 0.781 0.800 17.44 Wing Tail 0.205 0.250 4.46 Horizontal 0.125 0.149 3.25 Vertical 0.125 0.056 1.21 0.375 0.621 13.53 Landing Gear Main Gear 0.410 0.250 8.93 Nose Gear 0.125 0.211 4.60 Glue, etc. **** 0.200 4.35 Error Factor 5% 5% **** Total Unloaded 4.384 4.546 **** Payload 0.035 0.050 1.08 Total Loaded 4.419 4.596 **** A-10 Figure A-11: V-n Diagram 1.5

/

<-Vstal} =20.5 Vmax=,_3 -> 0.5

/

"O O -0.5

\

\

-1 0 10 2O 30 40 50 60 70 80 Velocity (feet per second) A-11 Figure A -12: Three View External Sketch Sw = 6.33 ft 2 Sv = .50 ft 2 Cw = .96 ft cv = .41 ft Sh = 1.50 ft 2 length = 3.54 ft bh = 2.74 ft Zinser 12 - 8 propeller Scale: I inch = 1.74 feet &--- \i

I

: I

I J A- 12 Figuxe A - 13: In_ _guration wing spars avionics batteries ,r_ _ii _ passengers

motor

speed controller servos batteries 4 channel receiver speed controller batteries wing spar

-oooo _A

motor __Wingtpar servos passengers 4 channel receiver and avionics batteries A-13

Appendix B: Critical Data Summary Parameter Initials of Date: 24 Mar RI: *[all distances relative Lo aircraft nose _nd in common units]* DESIGN GOALS: V cruise DK 55 ft/s DK 6 No. of passengers/crew DK 16000 ft Max Range at Wmax DK 5.5 lb Maximum TO Weight- WMTO BASIC CONFIG.

7r Wing Area 6.33 sq ft 4.6 lbs Maximum TO Weight - sg WMTO _mpty Flight Weight sg 4.54 lbs Wing loading(WMTO) sg 11.63 oz/sq ft r 40 in max length 6.75 ft max span r max height dk 24.6 in rotal Wetted Area dk 2048 sq in WING r 7.2 Aspect Ratio Span r 6.75 ft Area r 6.33 sq ft Root Chord r 0.937 ft rip Chord r 0.937 ft taper Ratio r 1 C mac - MAC r -0.12 r 0 leading edge Sweep 1/4 chord Sweep * Dihedral 5 deg r rwist (washout) Airfoil section r FX 63-137 Design Reynolds number r 328,000 t/c r 13.60% r Incidence angle (root) 0 deg B-1

Hor. pos of 1/4 MAC r 0.917ft

Ver. pos of 1/4 MAC r 0.71ft

e- Oswald efficiency r 0.8

CDo -wing r 0.007

CLo - wing r 0.456

r 4.35/rad

ZLalpha -wing

FUSELAGE

Length 40 in

Zross section shape square Nominal Cross Section Area 0.0625 sq ft Finess ratio 12.4 Payload volume 0.54 sq in Planform area 0.972 sq ft Frontal area 12.25 sq in ZDo - fuselage 0.005 EMPENNAGE Horizontal tail Area ]r 1.5 sq ft span ]r 2.74 ft ]r aspect ratio 0.548 ft root chord ]r tip chord ]r 0.548 ft average chord ir 0.548 ft taper ratio ir 1 l.e. sweep ]r 0 ]r 1/4 chord sweep ]r -2.25 deg incidence angle hor. pos. of 1/4 MAC ]r 2.958 ft ver. pos. of 1/4 MAC ]r 0.71 ft Airfoil section ]r NACA 0009 0.8 e - Oswald efficiency jr 0.0019 CDo -horizontal jr CLo-horizontal jr 4.46/rad jr CLalpha - horizontal 0.518 CLde - horizontal jr CM mac - horizontal jr Vertical Tail Area jr 0.5 sq ft

jr Aspect Ratio

0.41 ft root chord

jr

B-2

tip chord r 0.41ft

_veragechord r 0.41ft

Itaperratio r 1

r 0

I.e. sweep

r 0 il/4 chord sweep r 2.958 hor. pos. of 1/4 MAC 1.41 ft vert. pos. of 1/4 MAC Airfoil section NACA 0009 r SUMMARY AERODYNAMICS C1 max (airfoil) r 1.7 Cmo (airfoil) r -0.12 CL max (aircraft) r 1.46 lift curve slope (aircraft) r 4.30 /rad CDo (aircraft) r 0.0288 efficiency - e (aircraft) r 0.76 Alpha stall (aircraft) r 11 deg Alpha zero lift (aircraft) r -6 deg 12.22 L/D max (aircraft) iJ r 2.6 deg Alpha L/D max (aircraft) jr WEIGHTS 4.54 lbs Weight total (empty) sg 9.76 in C.G. most forward-x&y sg 10.03 in C.G. most aft- x&y

sg

Avionics 0.433 lb

sg

0.05 lb Payload-Pass.&lugg.-max sg 0.636 lb

sg

Engine & Engine Controls 0.1 lb Propeller

sg

1.056 lb Fuel (battery)

sg

1.906 lb Structure

sg

0.8 lb Wing

sg

0.485 lb Fusela_e/emp.

s8

0.621 lb

Landing gear sg

PROPULSION ASTRO 15 js Type of engines DK number 1 DK 5in placement Pavil max at cruise 155.93 Watts JS 39.385 Watts Preq cruise JS max. current draw at TO JS 16.7 amps B-3

cruisecurrent draw FS 7.05amps

FS

Propeller type Zinger

12 in

Propeller diameter DK

DK 8 in

Propeller pitch

Number of blades DK 2

S 7271.7

max.prop. rpm

4793.8 _S

cruiseprop. rpm

max.thrust S 3.12 lb

cruisethrust 0.528 lb

Is

P-100SCR battery type ]JS number !JS 1000 mah js individual capacity 1.2 V

js

individual voltage 1000 mah js pack capacity 15.6 V js pack voltage STAB AND CONTROL 0.492 c Neutral point jr 22.3 Static margin %MAC jr Hor. tail volume ratio 0.51 jr 0.17 Vert. tail volume ratio jr Elevator area jr 0.75 sq ft Elevator max deflection jr 30 deg Rudder Area jr I0.25 sq ft Rudder max deflection jr 30 deg Aileron Area l0

jr

Aileron max deflection 0 ir -0.971 ]r Cm alpha ,_n beta ]r 0.521 C1 alpha tail ir 0.518 C1 delta e tail ]r 0.319 PERFORMANCE Vmin at WMTO JS 20.45 ft/s Vmax at WMTO [S 84.6 ft/s Vstall at WMTO IS 20.45 ft/s Ran_:e max at WMTO 28692.2 ft

sg

'I' Endurance @ Rmax 10.63 rain

sg

Endurance Max at WMTO 13.92 rain

s_

25055.1 ft Range at @Emax sg 28842.7 ft Range max at Wmin

sg

ROC max at WMTO 20.0 ft / s sg rs 52177 ft Abs. Ceiling B-4

Min Glide angle JS 3.82degrees

LT/Odistance at WMTO IJS 21.3ft

SYSTEMS

DK tricycle

Landing gear type

DK 13.45

Main gear position

DK 7.62

Main gear length

DK 2.5 in dia

Main gear tire size

DK i4

nosegear position

DK 8

nosegear length

DK 2.5 in dia

nosegeartire size

js Itekin

enginespeed control

Control surfaces

jr rudder, elevator

ECONOMICS: ke

raw materials cost

ke 15179 ke $201

propulsion system cost

ke $234

_vionics system.cost

)roduction manhours ke 90 hours

_ersonnelcosts

ke $900 ke $100

Loolingcosts

ke $2562.11

Lotalcostper aircraft

B-5

Appendix C: Aircraft Data Base ib (ft) iPlane Name S (ft^2) Wto (Ib) Fus. shape Ifus (in) Wfus (oz) 4.38 5.84 2.75 Rec 17 10 FX/90, 1990 Blue Emu, 1993 5.6 Rec 58.8 10.44 10 10 60 13.76 10.94 8.75 iGold Rush, 1993 5.4 Rec 64 17 9.5 9.5 ]'he Airplane, 1993 5.25 Rec 42 9.6 4.67 7 ]'he Penguin, 1990 3.125 Rec, tapered 37 10.9 5.47 8 ]'he Screem-J4D, 1990 3 Rec, tapered 41 6 8.5 ]'he Drag-n-Fly 3.05 Rec, tapered 66 9.93 9.17 ]'he RTL-46, 1993 5.1 Rec 9.65 9.65 ]'he Diamondback, 1993 6.41 Rec 67 19.67 10 9.22 The Bunny, 1993 5.3 Trap 58 Plane Name Weight/S Win q Weight (Ib) Win£1 Loadin 9 (Ib/ft^2) 0.656 0.150 0.616 FX/90, 1990 Blue Emu, 1993 0.5625 1.96 0.196 Gold Rush, 1993 0.49 0.84 0.078 The Airplane, 1993 0.58125 1 0.105 The Penguin, 1990 0.66921 0.781 0.167 The Screem-J4D, 1990 0.548 0.648 0.118 The Drag-n-Fly 0.444 0.525 0.088 The RTL-46, 1993 0.514 1.31 0.131 The Diamondback, 1993 0.664 0.81 0.084 The Bunny, 1993 0.503 0.9 0.090 Clmax AR Plane Name lWei_ht/b Airfoil 1.23 7.79 FX/90, 1990 0.112 FX-63-137B-Pt 1.1 10 0.196 Wortman Blue Emu, 1993 1.6 7 Gold Rush, 1993 0.096 FX63-137 1.28 9.5 The Airplane, 1993 0.105 SPICA 1.1 10.5 The Penguin, 1990 0.112 Wortmann FX-63-137 1.3 11.7 The Screem-J4D, 1990 0.081 NACA 4415 1 12 The Drag-n-Fly 0.062 SPICA 1.8 8.46 The RTL-46, 1993 0.143 SD7062 1.17 9.65 The Diamondback, 1993 0.084 Clark-Y 1.45 8.5 0.098FX63-137 _he Bunny, 1993 C-1

PlaneName Ltver

Sv ibv (in) ARv Lt hor (ft)

High/Low,

Dihedral

(in^2) (ft)

50.46 8.4 2.05 1.4 2.05

FX/90,1990

H, 13deg

105.12 13.4 3 1.7 2.77

BlueEmu,1993

H,8 deg

144 12 1 4.475

Gold Rush, 1993 H, 15 deg 4.25

180 30 5 5.167

The Airplane, 1993 H, 8 deg 5.167

60.48 12 2.4 2.86

The Penguin, 1990 H, 3 deg 2.86

54.72 8.7 1.4 2.21

The Screem-J4D, 1990 L, 10 deg 2.06

72 12 2 2.3

The Drag-n-Fly H, 10.5 deg 2.3

105.12 15.21 2.2 3.4

The RTL-46, 1993 L, 10 deg 3.5 3.1

155.5 10.8

The Diamondback, 1993 J, 8 deg

169.92 18.43 4.57 2 4.4

The Bunny, 1993 L, 6 deg poly Plane Name ARh En£1ine, #, position Sh {in^2) bh (in) We (oz/ 69.4 13.88 2.78 3.5 FX/90, 1990 Astro 05 geared, 1, f 2.73 4.1 Blue Emu, 1993 221.76 24.6 Astro 15, 1, f 3.3 5.968 Gold Rush, 1993 230.4 27.6 Astro 25, 1, f 1.67 3.8 The Airplane, 1993 120 14.14 Astro 15, 1, f 4.16 Astro 15, 1, f 1.5 The Penguin, 1990 149.76 24.96 The Screem-J4D, 1990 90.72 15.65 2.72i Astro 05,1, f 3_ Astro 05, 1, f The Drag-n-Fly 144 22.4 3.26' The RTL-46, 1993 276.48 30 Astro 15, 1, f The Diamondback, 1993 1389.61 115.8 9.65 Astro 15, 1, f 429.12, 32.76 2.5 6.3 Astro 15, 1, f The Bunny, 1993 Plane Name iBatteries Wbatt (oz) Wencj (oz) Power (W) 6.5 115 5.46 FX/90, 1990 10.24 20O 13.53 Blue Emu, 1993 Ill P90SCR 900 13 P90SCR 900 16.9 Gold Rush, 1993 14.176 300 14.76 10.3, 2O0 12 P90SCR 900 The Airplane, 1993 10.3 200 The Penguin, 1990 7 AA 5.46 The Screem-J4D, 1990 6.5: 115 18 500MAH The Drag-n-Fly 200 12 POSC 900 The RTL-46,1993 200 12 Pg0SCR 9O0 The Diamondback, 1993 The Bunny, 1993 8.62! 200 13 P90-SCR 1000 18 C-2 Plane Name IProp type, # blades IProp diam./in I Wprop (oz / Prop pitch FX/90, 1990 10-6, 2 Blue Emu, 1993 Topfiite 12-6, 2 0.5 6 12 Gold Rush, 1993 IZinger J, 2 0.98 6 13 The Airplane, 1993 ,Zinger 1 8 12 The Penguin, 1990 'Zinger 10-4 2 4 10 1 6 10 _Tornado 10-6 The Screem-J4D, 1990 6 10 Zinger 10-6 The Drag-n-Fly The RTL-46, 1993 Zinger 12.5-6 6 12.5 The Diamondback, 1993 Zinger 11-7 7 11 0.69 6 The Bunny, 1993 Zinger 12-6 Plane Name Elevators?, Size, Def Ailerons?, Size, Def Flaps?, Size, Def N N FX/90,1990 Y, 5.97, 10 Blue Emu, 1993 N N Y, 9.8, 20 Gold Rush, 1993 N N Y, 115.2, 20 The Airplane, 1993 N N Y, 97.92, 16 The Penguin, 1990 Y, 80.7, N Y,52.5,+30/-20 The Screem-J4D, 1990 N N Y, 30.24, 25 ]-he Drag-n-Fly Y, Y, The RTL-46, 1993 N Y, 350, 20 Y,33.12, 15 The Diamondback, 1993 N N Y, 59.04, 10/-20 N Y, 85.7,18 Y, 100.7,20 The Bunny, 1993 Plane Name Rudder?,Size, Def CG {in/: fore, aft Vmax (ftJs) FX/90, 1990 Y, 29.78,20 Blue Emu, 1993 Y, 57.5, 20 16.6,18 51.3 Gold Rush, 1993 Y, 79.2, 45 17.46,18.396 49 The Airplane, 1993 Y, 72, 30 19.36, 20.2 54.3 The Penguin, 1990 Y, 42, 20 13.1,13.9 56.1 The Screem-J4D, 1990 Y, 30.24, 25 9.71 58 The Drag-n-Fly Y, 36, 20 9.4 The RTL-46, 1993 Y, 56.16, 30 17.7, 19 54 The Diamondback, 1993 Y, 97.92, 20 23.3 aft 59.5 5O The Bunny, 1993 21.23,23.3 Y, 102.24, 30 C-3 Plane Name i Vto (ft/s) Vstall (ft/s) Vcruise (ft/s) Range (ft): cruise, max 24 20.8 24 12210.,14389.

FX/90, 1990 Blue Emu, 1993 25.3 22 23169.,23667.

17.2_ Gold Rush, 1993 31 16600.,20250.

The Airplane, 1993 23 19.3 31 12100, 12500 22.6 25 2609, 20.

The Penguin, 1990 (23.7) 22.8 19 23 The Screem-J4D, 1990 5500, The Drag-n-Fly 25 4831, The RTL-46, 1993 23.28 19.4 35 19430 The Diamondback, 1993 21.7 17.8 28 18300 21.7 15.95 The Bunny, 1993 30 14325.3, 14728.

Plane Name Wav (oz) Endurance(min): Take-off Dist. (ft) Wg (oz) cruise,max 7.38 FX/90, 1990 7.99,8.48 30.96 3.5 5.92 Blue Emu, 1993 12.87,14.3 16.3 6 Gold Rush, 1993 10.3,10.3 24 5 5.95 The Airplane, 1993 6.8, 8.7 51.2 4 7.8 The Penguin, 1990 1.755, 45 2.4 4.15 The Screem-J4D, 1990 3.68, The Drag-n-Fly 3.22, 15.4 The RTL-46, 1993 13.52 25.4 The Diamondback, 1993 13.2 16.1 6 The Bunny, 1993 8.25, 10.25 (2-4

Appendix D: Performance Program This program, written by Sean Greenwood and Jeff Scherock, calculates the performance data for a given engine and propeller.

PROGRAM TS3 C REAL kt,kv, lf,J,Nprop,Nm,NmC,ia,NmCn C OPEN(UNIT=12,FILE='end') OPEN(UNIT=13,FILE='rang') OPEN(UNIT=14,FILE='sdrain') OPEN (UNIT = 15,FILE = 's time') OPEN(UNIT=16,FILE='Pre') OPEN(UNIT=IT, FILE='Pav') OPEN(UNIT=18,FILE='Roc') OPEN(UNIT=19,FILE='VH') OPEN(UNIT=20,FILE='par a') OPEN(UNIT =21,FILE ='ind') OPEN(UNIT=22,FILE='Treq') OPEN (UNIT=23,FILE ='Ppara') OPEN(UNIT=24,FILE='Pind') C Pi=4.*ATAN(1.0) C A1=0.3361362 A2=0.9234661 A3=0.6159474 A4=-1.352034 B1=0.01321402 B2--0.2063333 B3=-0.3656616 B4=0.1436228 C Aa1=-0.2729091 Aa2=2.906734 Aa3=-1.026263 Aa4=-1.56229 Bb1=-0.002688695 Bb2=0.04320921 Bb3---0.07096458 Bb4=0.0289369 C Ra=0.12 Rb=0.08 D-1 kv=0.000794 kt=1.097846 etag=.95 Tloss=l.372935 rho=0.002378 AR=7.2 b--6.75 S=b**2/AR cdo=0.021 If--1.

w=4.5 e=0.82 dprop=l.

grat=2.21 C Vact=9.6 CC V--50.

C CC DO 2 Range=10000.,20000.,2500.

DO 5 V=25.,90.,5.

C CI=lf*w*2*AR/(rho*V**2*b**2) Cd=cdo+(Cl**2/(Pi*e*AR)) Preq=1.356*rho*V**3*b**2*Cd/(2.*AR) C NmC--5000.

C 10 CONTINUE CONTINUE J=V*60*grat/(NmC*dprop) eta=AI+A2*J+A3*J**2+A4*J**3 Cp=(BI+B2*J+B3*J**2+B4*J**3) etaA=Aal +Aa2*J+Aa3*J**2+Aa4*J**3 CpA=2.*Pi*(Bbl+Bb2*J+Bb3*J**2+Bb4*J**3) Nprop=NmC/grat Pmob=1.356*Cp*rho*(Nprop/60.)**3*(dprop**5/etag) ia=(Pmob/(0.0007397*NmC*k0)+(Tloss/kt) Pmoa=0.0007397*NmC*(kt*ia-Tloss) Pavail=Pmoa*etag*eta ROC =0.7376 *(Pavail-Pre q) / w Nm=(Vact-ia*(Ra+Rb))/kv IF (ABS(Nm-NmC).GT.0.5) THEN NmCn=NmC+(Nm-NMc)/2.

NmC=NmCn GOTO 20 D-2 ENDIF IF (ABS(Pavail-Preq).GT.0.5) THEN Vactn=Vact-(Pavail-Preq)/Pavail Vact=Vactn GOTO 10 ENDIF C CCC WRITE(6,*)'eta,etaA',eta,etaA CCC WRITE(6,*)'Cp,CpA',Cp,CpA C Trp=cdo*.5*rho*V**2*S Tri=CL**2 / (Pi*AR*e)*.5*rho*S*V**2 Prp=Trp*V*1.356 Pri=Tri*V*1.356 Treq=Trp+Tri Time=970.*3.6/ia Range=V'Time angle=ASINCROC/V) VH=V*COS(angle) C WRITE(12,*)V,Time WRITE(13,*)V, Range CCC WRITE(20,*)V,Trp CCC WRITE(21,*)V,Tri CCC WRITE(22,*)V,Treq CCC WRITE(23,*)V,Prp CCC WRITE(24,*)V,Pri CCC WRITE(16,*)V,Preq CCC WRITE(17,*)V,Pavail CCC

WRITE(18,*)V_ROC

CCC WRITE(19,*)VH,ROC C CC Time=Range/V CC drain=ia*Time/3.6 CC WRITE(14,*)V, drain CC WRITE(15,*)V,Time C WRITE(6,*)' V,Vact',V, Vact wRrrE(6,*)' Nm,NmC,J',Nm,NmC,J WRITE(6,*)' ROC',ROC WRITE(6,*)' Range',Range WRITE(6,*)' Time',Time WRITE(6,*)' Pavail,Preq',Pavail,Preq WRITE(6,*)' Treqi,Treqp,Treq',Trp,Tri,Treq WRITE(6,*)' Pmoa,Pmob,ia',Pmoa,Pmob, ia WRITE(6,*)' ' D-3 C 5 CON'I'I]'qUE CC

WRITE(14,*)' '

CC

WRITEflS,*)' '

CC 2 CONTINUE C CLOSE(UNIT=12) CLOSE(UNIT=13) CLOSE(UNIT=14) CLOSE(UNIT=IS) CLOSE(UNIT=16) CLOSE(UNIT=17) CLOSE(UNIT=18) CLOSE(UNIT=19) CLOSE(UNIT=20) CLOSE(UNIT=21) CLOSE(UNIT=22) CLOSE(UNIT=23) CLOSE(UNIT=24) C END D-4

Appendix E: Aerodynamics, Stability, and Control Program program design implicit real(a-z) open(unit=100,file='data',status='unknown') C input constants pi=4.*atan(1.)

rhosl=0.0023769 gamma=l.4 p0=2116.2 grav=32.2 C input variables read(100,*) ew read(100,*) eac read(100,*) clas read(100,*) cmacw read(100,*) ast read(100,*) alo read(100,*) iw read(100,*) Gam read(100,*) ARw read(100,*) cf read(100,*) delf read(100,*) fpi read(100,*) fpo read(100,*) tauf read(100,*) delcdf read(100,*) kf read(100,*) read(100,*) Df read(100,*) Lf read(100,*) wf read(100,*) deuda read(100,*) k2kl read(100,*) Sv read(100,*) ARv read(100,*) try read(100,*) clasv read(100,*) kn read(100,*) krl read(100,*) Sh E-I read(10G*) ARh read(100,*) trh read(100,*) clash read(100,*) ih ih=ih*pi/180.

read(100,*) read(100,*) nult read(100,*) Wtot read(100,*) xcg read(100,*) ycg read(100,*) xw read(100,*) yw read(100,*) xh read(100,*) yh read(100,*) xv read(100,*) yv read(100,*) xmg read(100,*) read(100,*) trad read(100,*) tvel read(100,*) talp read(100,*) act read(100,*) vto read(100,*) ato read(100,*) vmax read(100,*) wv read(100,*) read(100,*) maxde read(100,*) taue read(100,*) maxdr read(100,*) taur read(100,*) maxda read(100,*) taua read(100,*) ya2 read(100,*) yal read(100,*) prat read(100,*) sigma close(unit=100) find n based upon performance requirements C n=sqrt(1. + (tvel**2 / gray / trad)**2) find wing cl E-2 clas=clas*180./pi clash=clash*180./pi clasv=clasv*180./pi claw=clas / (1.+clas / (pi*ew*ARw)) clat= clash / (1.+clash / (pi*ew*ARh)) clav=clasv/(1.+clasv/(pi*ew*ARv)) cla ac=cl as / (1. +clas / (pi*eac*ARw)) claacf=clas / (1 .+clas / (pi*eac*(1. +kf)*ARw)) CLmaxw=claw*(ast-alo+iw)*pi/180.

CLmaxac=claac*(ast-alo+(fpo-fpi)*tauf*d elf-2)*pi / 180.

find wing attributes based on performance CLper=claac*(talp-alo+iw)*pi/180.

Swl=n*Wtot/(CLper*0.5*rhosl*tvel**2) cwl=Swl/(sqrt(ARw*Swl)) bwl=sqrt(ARw*Swl) Gam=Gam*pi/180.

vstalll=sqrt(2.*Wtot/(rhosl*Swl*CLmaxac)) c find wing attributes based upon take-off CLto=Claacf*(iw+ato-alo+ (fpo- fpi)*tauf*delf)*pi / 180.

Sw2=Wtot/(CLto*0.5*rhosl*(vto)**2) cw2=Sw2 / (sqr t(ARw*Sw2)) bw2=sqrt(ARw*Sw2) vstall2=sqrt(2.*Wtot / (rhosl*Sw2*CLmaxac)) find wing attributes if (Swl.gt.Sw2) then Sw=Swl bw=bwl CW=CWl type=l else Sw=Sw2 bw=bw2 CW=CW2 type=2 endif find static stability coefficients cmaw=claw*(xcg-xw) / cw Vh=Sh*(xh-xcg) / (Sw*cw) E-3 deda=2*claw/(pi*ARw) e0=deda*(0.-alo)*pi/180.

cmat=-Vh*clat*(1-deda) cmaf= 1./(36.5*Sw*cw)*wf**2*deud a*Lf cma=cmaw+cmat+cmaf cm0f=k2kl / (36.5*Sw*cw)*wf**2*alo*pi/180.*Lf cm0w=cmacw+claw*(0.-alo)*pi/180.*(xcg-xw)/cw cm0t=vh*clat*(e0+iw*pi/180.-ih) cm0=cm0f+cm0w+cm0t Vv=Sv*(xv-xcg)/(Sw*bw) zw=Df/2.

dsdb=0.724+3.06*(Sv/Sw)/2.+0.4*zw/Df+0.009*ARv cnbv=Vv*clav*(dsdb) sfs=wf*Lf cnbwf=-kn*krl*sfs/Sw*Lf/bw cnb=cnbv+cnbwf clb =-2.*Gam*cw*claw*bw**2 / (Sw*bw*8) find drag coefficients cd0=(0.007*Sw+0.11*Df**2+0.01*Sw+0.008*(Sv+Sh))/Sw cd0=cd0*1.15 k=l./(pi*eac*ARw) C find cruise conditions vcrl=sqrt(wtot/(0.5*rhosl*Sw*claac*(iw-alo)*pi/180.)) vcr2 =sqrt(wtot / (0.5*rhosl*Sw*claac*(a cr+ iw-alo)*pi / 180. )) alpcr=2.*wtot/(rhosl*Sw*claac*pi / 180.*vvv**2)-iw+alo CLcrac1=claac*(iw-alo)*pi/180.

CLcrac2=claac*(acr+iw-alo)*pi/180.

CLcrac3=claac*(alpcr+iw-alo)*pi/180.

cdcrl=cd0+k*CLcracl**2 cdcr2=cd0+k*CLcrac2**2 cdcr3=cd0+k*CLcrac3**2 Dcr=cdcr*0.5*rhosl*vcr**2*Sw Treq=Dcr Mcr=vcr/sqrt(gamma*p0*prat/(rhosl*sigma)) LDcrl=CLcracl/cdcrl LDcr2=CLcrac2 / cdcr2 LDcr3=CLcrac3/cdcr3 open(unit=150,file='ld.out') do 10 v=25.,60.,2.5 ccll=wtot*2. / (rhosl*v**2*Sw) E-4 ccdd=cdO+k*ccll**2 lldd=ccll/ccdd write(150,*) v,lldd continue close(unit=150) find static margin to check again for pitch stability xnpalt=0.25-(cmaf+cmat) / claw xcgalt=(xcg-xw)/cw+0.25 sm=xnpalt-xcgalt find coefficients for static control and compare to requirements C clde=Sh / Sw*clat*taue cmde=-(Vh*clat*taue) dcmmax=cmde*maxde/180.*pi cmmax=cm0+cma*(ast)*pi/180.

cndr=-(Vv*clav*taur) dcnmax=cndr*maxdr / 180.*pi cnmax=cnb*15.*pi / 180.

clda=2.*taua*claw/(Sw*bw)*cw*0.5*((bw*ya2)**2-(bw*yal)**2) dclmax=clda*maxda/180.*pi cldr=Sv/Sw/bw*(0.5*sqrt(sv*arv))*taur*clav takeoff performance liftt=clat*ih*0.5*rhosl*(0.85*vto)**2*Sh-clde*maxde/180.*pi liftw=claw*(0.-alo+ (fpo-fpi)*tau f*delf)*pi/180.*0.5*rhosl & *(0.85*vto)**2*Sw mtail=-liftt*(xh-xmg) mwing=-liftw*(xw-xmg) mweight=wtot*(xcg-xmg) mcg=mtail+mwing+mweight C performance characteristics phi= (16./pi*(yw-ymg)/bw)**2 / (1+(16./pi*(yw-ymg)/bw)**2) cdto=cd0+k*clto**2*phi+0.02 cdtoa=cdto-0.02 dto=0.5*rhosl*vto**2*Sw*(cdto) dtoa=dto/cdto*cdtoa minr=vcr**2 / (32.2*sqrt(nult**2-1 .)) CLm=sqrt(cd0/k) CDm=cd0+k*CLm**2 E-5

LDmax=CLm/CDm

deltcr=-(cm0*CLaac+cma*CLcrac3) / (cmd e*CLa ac-cma*clde)

output write(6,*) write(6,*) 'weight',Wtot write(6,*) 'Wing loading',Wtot*16./Sw ! i write(6,*) xcg,ycg ,xcg,ycg write(6,*) 'n during tum',n I t write(6,*) write(6,*) 'Wing area',Sw,type write(G*) 'Wing span',bw write(G*) 'Wing chord',cw 'AR',ARw write(G*) 'eac',eac write(G*) p, , • write(G*) 1W ,lW ! !

cf ,cf write(6,*) write(G*) 'inner and outer flap position',fpi,fpo write(6,*) 'Swl,Sw2',Swl,Sw2 v write(6,*) write(6,*) 'max lift angle',ast write(G*) 'zero lift angle',alo write(6,*) 'cruise angle',acr write(6,*) 'take-off angle',ato write(G*) 'tum angle',talp write(6,*) 'max flap deflection',delf ! I write(6,*) 'CLmax aircraft',CLmaxac write(6,*) write(6,*) 'CLto',CLto write(6,*) 'CLper',CLper write(6,*) 'CLalpha aircraft',CLaac write(6,*) 'CLaplha wing',claw 'CLcruise aircraft',CLcracl,CLcrac2,CLcrac3 write(6,*) 'CDcruise aircraft',cdcrl,cdcr2,cdcr3 write(G*) 'cd0,k',cd0,k write(6,*) 'cdto,cdtoa',cdto,cdtoa write(G*) 'dto,dtoa',dto,dtoa write(6,*) t !

write(6,*) 'Vto',vto,vto/vstall2 write(6,*) 'Vstall',vstalll,vstall2 write(6,*) write(G*) 'Vcruise at 0 degrees',vcrl write(6,*) 'Vcruise at ',acr,' degrees',vcr2 write(G*) 'Alpha cruise at ',vvv,' ft/s',alpcr write(6,*) 'Mcr',Mcr E-6 write(6,*) 'max vel',vmax write(G*) 'Minimum turn radius at cruise',minr write(G*) 'L/D cruise',LDcrl,LDcr2,LDcr3 write(G*) 'L/D max,CLm',LDmax,CLm g !

write(6,*) write(G*) 'Sv guess, Cnb',Sv,Cnb write(6,*) 'Sh guess, Cma, Cm0',Sh,Cma,Cm0 write(6,*) 'dihedral guess, Clb',Gam*180./pi,Clb 'Vh,Vv',Vh,Vv write(6,*) 'sm',sm write(G*) 'dcmmax',dcmmax write(G*) 'cmmax',cmmax write(G*) 'clde',clde write(6,*) write(6,*) 'cmde',cmde write(G*) 'dcnmax',dcnmax write(6,*) 'cnmax',cnmax write(6,*) 'cndr',cndr write(6,*) 'dclmax',dclmax 'clb',clb write(G*) write(6,*) 'cldr',cldr write(6,*) 'delta trim cruise',deltcr*180./pi write(6,*) 'xloc',xw+0.5*cw write(G*) write(G*) 'horiz xloc',xh+0.5*(Sh/sqrt(Sh*ARh)) write(6,*) 'vert xloc',xv+0.5*(Sv/sqrt(Sv*ARv)) write(G*) 'xnpalt',xnpalt write(6,*) 'xcgalt',xcgalt write(G*) 'Mcg',mcg write(6,*) open(unit=159,file='coeffs.dat') write(159,*) cm0f, cmaf write(159,*) cm0w,cmaw write(159,*) cm0t, cmat write(159,*) cnbwf, cnbv close(unit=159) stop end E-7

Appendix F: Manufacturing Plan

F.1 Overview Long Shot Aeronautics is faced with the difficult task of assembling a prototype of The Balsa Bullet. The single largest hurdle to be overcome in this task is experience. The group members of Long Shot Aeronautics hold no experience in any RPV manufacturing techniques. For this reason, this manufacturing plan is set forth to be a comprehensive guide to 1) get all group members "on the same page," and 2) insure that from our first assembling steps to our "coup de grace," none of our steps are timid. This plan will first identify major components and subsystems, after which the timetable will be set forth.

Following these will be a tooling schedule, critical methods, and will conclude with a cost analysis.

F.2 Major Components and Subsystems 1) Wing A) Main Spar B) Flaps C) Main Landing Gear 2) Fuselage A) Propulsion Mounts i) Engine Bulkhead ii) Battery Mount iii) Cowl B) Control Systems Mounts i) Servos ii) Avionics F-1

C) Nosewheel Mount

3) Empennage

A) Horizontal Tail/Elevator B) Vertical Tail/Rudder F.3 Timetable The timetable for manufacturing appears on the following page.

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F.4 Tooling Schedule A schedule and tracking sheet for tooling requirements appears on the following page.

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o tt3 t'N _°___ ° O F.5 Critical Methods F.5.1 Structural Assembly Techniques Without a doubt, the single most critical component of the aircraft assembly is the main spar joint. From consultation with Mr. Doug Staudmeister, the joint will be a butt joint reinforced with nylon mesh, carbon filament, and epoxy resin (figure F-l). The mesh and epoxy will also be used to fasten the main spar assembly to the fuselage longerons in the areas shown in figure F-1.

Fuselage Cross Section Spruce 1/4" X 1/4" Longeron v ¢#I ; ; - Spar Notched with Fuselage Brace Piece "" •-, Fuselage Braces ¢J, fJJ cJl fJJ _r #,-# Win Main Spar _. #,-# !#/._ 5 deg Plywood Sheets (Sandwiching Main Spar) Hardwood Blocks, Fastened to Spar and Plywood Plates Figure F-l: Wing Carry-Through Design F-4

Another critical component will be the main landing gear. Figure F-2

shows the configuration and attachment to the main spar. The complexity of the

bends will require patience,however, a slip in this processis correctable, yet

costly in time.

Rib Section

5/32 "Steel Rod

Foam Tire

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Birch Plywood Flooring Lower Spar Cap Figure F-2: Landing Gear Detail F.5.2 Personnel Management The philosophy behind personnel management is based on "relative expertise." Initial task assignments were distributed randomly, as no one person is qualified in any area, however, as time progresses, it is assumed that each F-5

member will become acquainted with a certain task. For this reason, two

members have been assigned construction duties for both the first and second wings, with the thought in mind that experience gained through manufacture of the first will be of use in fabricating the second. Similarly, two people will be assigned as machine operators for the first two days of construction. It is hoped that they will become proficient (in a relative way) on specific tools, saving on run time costs. The remaining two members will begin construction on the fuselage section.

As time progresses and machine time lags, the two machine operators will likely be reassigned the subsystem tasks of flaps and main landing gear. As the fuselage and empennage sections near completion, one of the responsible individuals will begin monokote testing and training, while the other, specifically the Director of Manufacturing, will provide assistance wherever it may be needed.

A rough budgeting of the ninety manufacturing hours is as follows: Wing 30 hours Fuselage 8 hours Empennage 9 hours Landing Gear (main) 7 hours Final Assembly 10 hours Installations 8 hours Monokoting 8 hours Miscellaneous 10 hours The wing construction for component validation is budgeted an additional 35 hours, since this will be the first wing built.

F-6 F.6 Cost Evaluation F.6.1 Raw Materials The raw material expenditures as of the Manufacturing Plan Review was broken down as follows: Spruce $21.52 Balsa $26.66 Plywood $15.42 Monokote $18.98 Landing Gear $23.17 Glue $ 7.18 Miscellaneous $24.66 Subtotal $137.59 Tax (5%) $ 6.88 TOTAL $144.47 F.6.2 Remaining Estimated Costs The remaining costs are estimated as follows: Fixed Subsystems $435.00 Labor $900.00 Tooling $ 43.85 Wa_t_ $ 80.00 Subtotal $1458.85 Raw Materials $ 144.47 New Estimate $1603.32 Qld Estimat¢ $1684.00 Surplus $ 81.68 F-7

Although arriving at the final product under budget would be a good

thing for The BalsaBullet, a preliminary qualitative analysis would lead Long

Shot Aeronautics to reallocate the surplus towards labor costs. Although a

preliminary labor schedulehas been laid out, totaling 90hours, difficulties are

anticipated particularly in fabricating the wing carry-through structure.

F-8

Appendix G:

Flight Validation, Component Test

and

Manufacturing Hours

p_mm,_ PAG_ ImmmLAr,_ NOT i_.M_ appeared to handle much better at low speed (about I/3 throttle setting). Turning performance was acceptable and much improved with the sealed rudder gap but there did not appear to be enough wing dihedral. A high speed leg was attempted and the speed estimated at 60 ft/sec - the turn from the high speed leg was difficult to achieve in the confines of Loftus. Take-off tests with the flaps deployed were not conducted.

Wing Component Static Load Test, April 19, 1994 Spring 1994 Balsa Bullet Summary; A wing component was tested to failure. The wing was completed with flaps and attached to a mockup of the center fuselage section.

The center fuselage section was constructed in a manner similar to the planned fuselage construction and then mounted to the static load facility with clamps. The wing weight including the fuselage centerbody was 0.81 Ib as tested.

The wing failed when a total load of 9.6 Ib was applied. Failure was due to a debonding of the adhesives at the wing centerline which were used to attach fiberglass "material" which was the only connection between the spar caps at the root. The caps were simply "butt-joined" at the root and wrapped with the fiberglass cloth. There did not appear to be any structural failure in the wings themselves.

Prior to failure either "cracking" or monokote debonding from the structure was "heard" at locations outboard from the root.

I-_ Load Distribution: The approximation to the 1-g load was applied starting at the root.

The spanwise locations where the loads were applied started 3" from the root and were spaced at 6" intervals. The l-g load was applied first and then the 2-g condition was applied by increasing the load starting at the root. This processes continued until the wing failed.

The wing failed when the total load applied to the wing was 9.6 Ib, and this occurred as the loading was being increased from 2 to 3 g's.

Spanwise location (distance from Load (Ib) root in inche s) 3 .39 9 .39 15 .39 21 .35 27 .35 33 .29 39 .12 Win_ Tip Deflection: The tip deflection was measured as the load was increased. It is presented for even increments in load factor and the last data point taken before failure.

Total Load (Ib) - Both wings Tip Deflection (in) 4.6 1.0 9.2 2.5 9.6 failure Additional Information: Aircraft Weight = 4.6 Ib (estimate at this time) Wing Weight = 0.8 1 Ib (as tested)

Comparison Between Design and Actual Aircraft Data

Actual Value

Design Value

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Wing Area _-._ ._._,= o_.:,.,_, ,,._,,:-

Vertical Tail Area

Horizontal Tail Area

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Wing Structural Weight (Monokote)

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Wing Structural Weight (no Monokote)

Fuselage Structural Weight (Monokote

Fuselage Structural Weight 'no Monokote)

led _ 0,_432 I;s

Vertical Tail Weight (Monokote)

Vertical Tail Weight (no Monokote)

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Horizontal Tail Weight (Monokote)

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Horizontal Tail Weight no Monokote)

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Landing Gear Weight

Propeller Type

Propeller Weight

Total Aircraft Weight (post-construction)

Total Aircraft Weight (post-flight)

CG Location (post-construction)

CG Location post-flight)

Weight of Batteries

'/79< I, c>m_:i,j Please list any other deviations of the technology demonstrater from the original design.

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Long Shot Aeronautics: Time Sheet I

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Long Shot Aeronautics: Tooling Sheet I

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