Canard Aircraft Design Project

Canard Aircraft Design Project

(Parte 1 de 2)

Aircraft Performance Dynamics and Design

Dr. Vijay Chatoorgoon December 6, 2013

MECH 4452 University of Manitoba

List of Figuresi
List of Tablesiv
Executive Summaryv
1 Introduction1
1.1 Background1
2 Design Constraints2
3 Design Methodology3
4 Design Details4
4.1 Airfoil Choice4
4.2 Dimension Choices6
4.3 Engine and Propeller12
4.4 Design Calculations14
5 Conclusion26
References27

Table of Contents Appendix ................................ ................................ ................................ ................................ ............... 28

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Figure 1. Moment Coefficient about the CG vs. Absolute Angle of Attack8
Figure 2. Aircraft Drawing 1 of 210
Figure 3. Aircraft Drawing 2 of 21
Figure 4. O.S. 61FX Engine12

List of Figures Figure 5. APC LP12060 Thrust Plot (10,0 RPM) ................................................................... 13

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TABLE I: MAIN WING CHARACTERISTICS (NACA 23012)5
TABLE I: CANARD WING CHARACTERISTICS (EPPLER EA 6(-1)-012)6
TABLE I: AIRCRAFT DIMENSIONS7
TABLE IV: STABILITY PARAMETERS (AT A CRUISE ANGLE OF 1.01225 DEGREES)8

List of Tables TABLE V: DENSITY VS. VELOCITY PLOT ......................................................................... 2

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Executive Summary

The purpose of this project was to design an aircraft in terms of aerodynamics, stability, and performance in compliance with the 2013 Aero Design Rules. An additional set of design constraints were also given describing that the aircraft had to be of ‘canard’ type optimized to lift maximum load. The resulting design had to provide a list of deliverables namely the aircraft dimensions, the maximum gross weight, the take-off and landing velocities, and the canard design.

The scope of this project chiefly focused on evaluating the performance of the aircraft design. Material selection, flow and stress analysis, and aircraft control were not considered in this analysis.

The approach used to design the aircraft involved first conducting an external search on airfoil selection, canard design, and propulsion design with the goal to maximize lift. The resulting airfoils and wing designs were iteratively compared in a program written to determine aircraft stability. With a stable design, the aircraft dimensions were fed into a second program written to evaluate the performance of the aircraft.

The final design sports a cambered semi-symmetric NACA 23012 airfoil for the main wing and a symmetric Eppler EA 6(-1)-012 for the horizontal and vertical stabilizers. The selected power plant mounted in a pushing configuration is an O.S. 61FX engine with a 12x6P APC pusher propeller developing 6.01 lb of thrust in static conditions. The design also features rectangular wing spoilers for landing purposes. Given the lack of material selection, the aircraft was assumed to weigh 8 lb empty.

The design was proven to be stable throughout its range of motion while maintaining maximum use of its size constraints. This design is expected to lift a payload of 13.5 lb while complying with the respective 200 and 400 ft take-off and landing limits. The maximum speed for the design is 103.29 ft/s with a maximum climb angle of 6.6°.

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1 Introduction

The design team was tasked to aerodynamically design an aircraft which was to conform to the Regular Class constraints of the 2013 SAE Aero Design Rules. The aircraft was also to conform to constraints set by Dr. Chatoorgoon. The aforementioned design constraints are outlined in full detail in the Design Constraints section of this report.

The aircraft’s mission is to lift the maximum load possible while conforming to the design constraints. The aircraft required a canard configuration where the tail-plane is located in front of the main wing. Some background for the canard configuration is given in the following subsection.

1.1 Background

A canard aircraft configuration features two lifting surfaces. However, unlike a conventional configuration, the horizontal stabilizer is located in front of the main wing; this is known as a canard configuration. A canard generates positive lift (up-lift) in order to balance the forces and moments about the center of gravity of the aircraft. The positive lift is beneficial for lifting more weight. Although, the downwash effect of the canard on the main wing reduces the lift on the main wing. In addition, according to [1], the canard must stall before the main wing. This means that using the maximum lift of the main wing at a maximum lift-angle over that of the canard can risk stalling the canard.

A canard configuration aircraft must have its center of gravity ahead of the aerodynamic center of the main wing to achieve static pitch stability, as determined by our analysis.

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2 Design Constraints

The team used the Regular Class constraints of the 2013 SAE Aero Design Rules for our design. The following related design constraints are outlined in the rules:

Takeoff distance limit: 200 ft Landing distance limit: 400 ft

Max combined length, width, height of 225 in (where length is the distance from front to aft, width is the span from wingtip to wingtip, and height is the distance from ground to highest point excluding propeller)

65 lb max gross weight (with payload & fuel)

Engine must be O.S. 61FX with E-4010 Muffler

Propeller must rotate at engine RPM

Addition design constraints were given by Dr. Chatoorgoon:

The aircraft configuration must be a canard type (tail-plane in the front) Ground temperature of 35oC

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3 Design Methodology

In order to facilitate the design, as well as provide values for some unknown parameters, several assumptions were made. The following list outlines all of the assumptions applicable to the design:

The zero-lift drag coefficient for the aircraft body itself was taken to be 0.005 in order to get the an approximation to drag created by the body of the aircraft

The thrust variation with aircraft speed was taken into account due to the aircraft being propeller driven. In order to facilitate calculations, the variation was taken to be linear The weight of the aircraft was assumed to be fully carried by the main wing

The downwash effect from the canard on the main wing was taken to be negligible in order to facilitate the analysis

The canard is designed to stall after the main wing rather than before due to a lack of information during the design finalization The empty weight of the aircraft is assumed to be 8 lb

It was assumed that 1.34 HP demand at 10,0 RPM can be satisfied by the engine

Standard atmospheric conditions were assumed, aside from the 35oC at sea level given by

Dr. Chatoorgoon The take-off and landing friction coefficients were assumed to be 0.04 for pavement

The main wing zero-lift line is parallel to the ground at take-off

 The spoiler drag coefficient is assumed to be a flat plate perpendicular to the air flow

(cd = 1.0) The spoiler is assumed to eliminate lift from the main wing

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4 Design Details

This section explains in detail the approach to the aircraft design. The first step was airfoil selection in section 4.1. The subsequent steps include dimensional and stability analysis, propulsion analysis, and performance analysis, found in sections 4.2, 4.3, and 4.4 respectively.

4.1 Airfoil Choice

It was proven in the derivation found in the appendix that the center of gravity of a canard configuration aircraft must be ahead of the aerodynamic center of the main wing for stability. The team concluded that to have the aircraft balanced and stable, the lift that is being generated by the canard must be positive.

The team developed a program in Microsoft Excel to simulate different parameters and dimensions aiming to get an optimum design for our aircraft. The tables in this program are included at the end of the appendix. After several attempts to get the aircraft stable, it was decided that high camber high-lift wing was unfeasible. Highly cambered airfoils have a very high pitching moment coefficient about its aerodynamic center. This high pitching moment did not allow for a positive CM,0, required for stability.

Due the problems encountered with high camber wings, a lower cambered wing with a higher stall angle was selected for analysis. The team decided on this because even though the wing does not have a high camber, due to the high stall angle and lower drag, a high lift wing with a low pitching moment about the aerodynamic center could be obtained. The resulting aircraft would be stable with a reasonable lift coefficient. For the main wing, the team decided upon using a NACA 23012, which gives an absolute maximum angle of attack of 16.75° and a CL,max of 1.5413. All the parameters of the main wing are shown in TABLE I.

The main wings also incorporate two extendable spoilers which act as air brakes for landing. In addition, they remove the lift from the wing by opening a gap between the top and bottom of the wing to eliminate the pressure difference. They are rectangular shaped with their surfaces extended (approximately) perpendicular to the air flow. At this position, they offer a drag coefficient of 1.0. Each wing also sports an aileron to provide rolling motion.

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The horizontal stabilizers were decided to function upon pivoting the entire canard. This eliminated the need to add elevator servos which reduces the weight of the aircraft. This design while being simpler however does not offer stick-free configuration for pitch control.

The vertical stabilizer shared the same airfoil with the horizontal stabilizers. The vertical stabilizer however was placed just in-front of the engine with the tip extending up to the height of the propeller.

It should be noted that control in the yaw and roll motions are not in the scope of this analysis. Rather, the vertical stabilizer and ailerons are mentioned to provide placement should an extended analysis be done.

TABLE I: MAIN WING CHARACTERISTICS (NACA 23012)

Oswald Efficiency 0.7 Infinite Wing Slope 0.0894 Finite Wing Slope 0.064626 degrees-1 Wing Chord 0.4318 meters Wing Span 2.62382 meters Aspect Ratio 6.076471 Wing Area 1.132965 meters² Pitching Moment Coefficient about the Aerodynamic Center -0.0067 Aerodynamic Center Position about the Leading Edge 0.25 % Maximum Absolute Angle of Attack 16.75 degrees

As stated in lecture, the canard wing must have as high a stall angle as possible, and has to be higher than the stall angle of the main wing for aircraft stability and stall recovery. Due to this fact, the team decided upon using an Eppler EA 6(-1)-012 for our canard wing because it has a high stall angle of attack of 18.25°, and could therefore provide a good margin to work on the stability design. All the parameters of the canard wing are shown in TABLE I.

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TABLE I: CANARD WING CHARACTERISTICS (EPPLER EA 6(-1)-012)

Wing Efficiency 0.9 Infinite Wing Slope 0.079616 Finite Wing Slope 0.060802 degrees-1 Wing Chord 0.3556 meters Wing Span 1.8542 meters Aspect Ratio 5.214286 Wing Area 0.659354 meters² Aerodynamic Center Position about the Leading Edge 0.25 % Maximum Absolute Angle of Attack 18.25 degrees-1

4.2 Dimension Choices

The first concern in the selection of the aircraft dimensions was its ability to conform to SAE rules. According to the rules, the sum of the height, length, and width of the aircraft must be less or equal to 225 in. Dimensional optimization was done over an iterative process aiming to have a high lifting main wing.

Another important point taken into account for setting our dimensions was that the stability conditions must be satisfied to ensure the aircraft stable and balanced. The dimensions of the main and canard wings have a strong impact on the stability of the aircraft, since the forces and moments generated for both wings have to be in balance about the center of gravity of the aircraft. Another important parameter in the calculations was the “tail volume” which has a strong influence on the length of the aircraft. All the dimension of our aircraft are shown in TABLE I.

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Angle of attack of the Canard Wing About the Zero Lift Line of the

Airplane 1 degrees

Tail Volume 1.48 Space for the Engine 0.0596 meters Airplane Height 0.3556 meters Total Length of the Airplane 2.721456036 meters Distance Between the Aerodynamics Centers 2.249106036 meters

Distance from the Aerodynamic Center of the Canard to the Center of

Gravity 1.098102045 meters

Distance from the Aerodynamic Center of the Main Wing to the Center of

Gravity 1.15100399 meters

Distance from the Aerodynamic Center of the Tail to the front of the

Aircraft 0.0889 meters

Distance from the Trailing Edge of the Tail to the front of the Aircraft 0.3556 meters Distance from the Center of Gravity to the front of the Aircraft 1.187002045 meters Distance from the Neutral Point to the front of the Aircraft 1.736753659 meters Distance Between the Center of Gravity and the Neutral Point 0.549751613 meters

Distance from the Leading Edge of the Main Wing to the front of the

Aircraft 2.230056036 meters

Distance from the Aerodynamic Center of the Main Wing to the front of the Aircraft 2.338006036 meters

Distance from the Trailing Edge of the Main Wing to the front of the

Aircraft 2.661856036 meters

Airplane Length + Wing Span + Height 5.700876036 meters

The stability parameters listed in TABLE IV show that all the stability parameters were satisfied for the aircraft at cruise conditions. It can be seen that the pitching moment coefficient when lift is zero, CM.0, is larger than zero. It can also be seen that the slope of the total pitching moment coefficient curve is negative. Furthermore, the static margin of our aircraft is shown, which defines how maneuverable it is. The higher this value, the better is the aircraft maneuverability.

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Finally, it is also shown that for cruise conditions, the total pitching moment coefficient of the aircraft is zero, which means our aircraft is in balance.

TABLE IV: STABILITY PARAMETERS (AT A CRUISE ANGLE OF 1.01225 DEGREES)

Pitching Moment Coefficient when the Lift is Zero (CM0) 0.083287118 Slope of the Total Pitching Moment Coefficient Curve -0.082279198 Static margin 1.273162606 CMcg 0

Figure 1 plots the total pitching moment coefficient about the center of gravity versus the absolute angle of attack of the aircraft. It can be seen again that we have a stable aircraft.

Figure 1. Total Pitching Moment Coefficient about the CG vs. Absolute Angle of Attack

The height of the aircraft was considered to be the distance from the ground to the tip of the vertical stabilizer. This dimension was fixed at 14 in where the 12 in propeller is located with its axis aligned with the zero-lift line of the main wing; midway between the top and bottom of the body. This figure allowed for a 2 in clearance between the ground to the propeller so that at rotation, the blades will not hit the ground.

T o t al Pitc h i ng

M o m e nt C o e f f i c i e nt a b o ut h e C

Absolute Angle of Attack [degrees]

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The following pages include engineering drawings of the calculated aircraft dimensions including the exact airfoil shapes in AutoCAD. Figure 2 and Figure 3 display a scaled side-view and top-view of the aircraft.

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10 Figure 2. Aircraft Drawing 1 of 2

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1 Figure 3. Aircraft Drawing 2 of 2

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4.3 Engine and Propeller

The engine specified by the SAE Aero Rules 2013 is the O.S. 61FX with E-4010 Muffler. The unit produces 1.9 BHP at 16,0 RPM. The technical specifications are as follows [2]:

(Parte 1 de 2)

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