INTRODUCTION

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Final Report (Volume II)

Raven Design Team

Brian HucksGroup LeaderRachel BlockStructuresRyan CzerwiecAerodynamics� XE "Aerodynamics" �Elton FairlessInternal Systems & Procurement� XE "Procurement" �Dustin KeckReport OrganizerRyan McDanielPerformance� XE "Performance" �Matt ShaverStability� XE "Stability" � & ControlsMichael SirmansWeights� XE "Weight" � & BalancesSean ZamoraCAD� XE "AutoCAD" � & Unigraphics� XE "Unigraphics" �

AE Senior Design Project (Group II)Advisor: Dr. John N. PerkinsDepartment of Mechanical & Aerospace EngineeringNCSURaleigh, NC5 May 1997�ABSTRACT

The need for short-range� XE "Range" � reconnaissance missions was the impetus for a tube-launched reconnaissance aircraft. The Raven, a ring-tailed Remotely Piloted Vehicle (RPV), was designed, constructed, and flight tested� XE "Flight test" � as a predecessor to this tube-launched solution. Issues of control surface configuration� XE "Configuration" � and sizing were central to the success of this project. Specifically, the ring-tail sizing was integral to stability� XE "Stability" �. Also, issues of yaw� XE "Yaw" � control were addressed extensively. Molded composite construction methods were utilized on the fuselage� XE "Fuselage" � and spars� XE "Spars" � while foam� XE "Foam" � core and balsa� XE "Balsa" � sheeting� XE "Sheeting" � techniques were used on the wing� XE "Wing" � and tail. The Raven aircraft was constructed on time and under budget� XE "Budget" �, with the first flight attempt on 2 April 1997. The first successful flight occurred on 24 April 1997.�Table of Contents� TOC \o "1-3" �1. Introduction� GOTOBUTTON _Toc387415157 � PAGEREF _Toc387415157 �7��2. Design History� GOTOBUTTON _Toc387415158 � PAGEREF _Toc387415158 �8��2.1 Wing� GOTOBUTTON _Toc387415159 � PAGEREF _Toc387415159 �8��2.2 Tail� GOTOBUTTON _Toc387415160 � PAGEREF _Toc387415160 �9��2.3 Skids� GOTOBUTTON _Toc387415161 � PAGEREF _Toc387415161 �9��3. Wind Tunnel Testing� GOTOBUTTON _Toc387415162 � PAGEREF _Toc387415162 �10��4. Stability and Controls� GOTOBUTTON _Toc387415163 � PAGEREF _Toc387415163 �11��4.1 Drag Rudder to Deflectors� GOTOBUTTON _Toc387415164 � PAGEREF _Toc387415164 �11��4.2 Controller� GOTOBUTTON _Toc387415165 � PAGEREF _Toc387415165 �11��4.3 Trim� GOTOBUTTON _Toc387415166 � PAGEREF _Toc387415166 �12��5. Structures� GOTOBUTTON _Toc387415167 � PAGEREF _Toc387415167 �13��5.1 Structures Review� GOTOBUTTON _Toc387415168 � PAGEREF _Toc387415168 �13��5.2 Finite Element Analysis� GOTOBUTTON _Toc387415169 � PAGEREF _Toc387415169 �13��5.3 Conclusions� GOTOBUTTON _Toc387415170 � PAGEREF _Toc387415170 �15��6. Weights and Balances� GOTOBUTTON _Toc387415171 � PAGEREF _Toc387415171 �16��7. Performance� GOTOBUTTON _Toc387415172 � PAGEREF _Toc387415172 �17��8. Construction� GOTOBUTTON _Toc387415173 � PAGEREF _Toc387415173 �18��8.1 Fuselage� GOTOBUTTON _Toc387415174 � PAGEREF _Toc387415174 �18��8.1.1 Plug� GOTOBUTTON _Toc387415175 � PAGEREF _Toc387415175 �18��8.2 Wing Construction� GOTOBUTTON _Toc387415176 � PAGEREF _Toc387415176 �22��8.3 Tail� GOTOBUTTON _Toc387415177 � PAGEREF _Toc387415177 �24��8.4 Painting� GOTOBUTTON _Toc387415178 � PAGEREF _Toc387415178 �26��9. Flight Testing� GOTOBUTTON _Toc387415179 � PAGEREF _Toc387415179 �27��9.1 Power Up Testing� GOTOBUTTON _Toc387415180 � PAGEREF _Toc387415180 �27��9.2 First Flight Attempt� GOTOBUTTON _Toc387415181 � PAGEREF _Toc387415181 �27��9.3 Second Flight Attempt� GOTOBUTTON _Toc387415182 � PAGEREF _Toc387415182 �29��9.4 Third Flight Attempt� GOTOBUTTON _Toc387415183 � PAGEREF _Toc387415183 �29��9.5 Fourth Flight Attempt� GOTOBUTTON _Toc387415184 � PAGEREF _Toc387415184 �30��10. Modifications and Repairs� GOTOBUTTON _Toc387415185 � PAGEREF _Toc387415185 �31��10.1 First Flight Attempt� GOTOBUTTON _Toc387415186 � PAGEREF _Toc387415186 �31��10.2 Second Flight Attempt� GOTOBUTTON _Toc387415187 � PAGEREF _Toc387415187 �32��10.3 Future Modifications� GOTOBUTTON _Toc387415188 � PAGEREF _Toc387415188 �33��11. Budget Overview� GOTOBUTTON _Toc387415189 � PAGEREF _Toc387415189 �34��12. Concluding remarks� GOTOBUTTON _Toc387415190 � PAGEREF _Toc387415190 �35��12.1 Conclusions� GOTOBUTTON _Toc387415191 � PAGEREF _Toc387415191 �35��12.2 Future Research� GOTOBUTTON _Toc387415192 � PAGEREF _Toc387415192 �35��13. Acknowledgments� GOTOBUTTON _Toc387415193 � PAGEREF _Toc387415193 �36��14. References� GOTOBUTTON _Toc387415194 � PAGEREF _Toc387415194 �37��15. Appendix� GOTOBUTTON _Toc387415195 � PAGEREF _Toc387415195 �38��15.1 Program Trim� GOTOBUTTON _Toc387415196 � PAGEREF _Toc387415196 �38��16. Tables� GOTOBUTTON _Toc387415197 � PAGEREF _Toc387415197 �41��16.1 Introduction� GOTOBUTTON _Toc387415198 � PAGEREF _Toc387415198 �41��16.2 Design History� GOTOBUTTON _Toc387415199 � PAGEREF _Toc387415199 �41��16.3 Wind Tunnel Testing� GOTOBUTTON _Toc387415200 � PAGEREF _Toc387415200 �42��16.3.1 New Pitch Rate and Yaw Rate Coefficients� GOTOBUTTON _Toc387415201 � PAGEREF _Toc387415201 �43��16.4 Stability and Controls� GOTOBUTTON _Toc387415202 � PAGEREF _Toc387415202 �44��16.4.1 Stability Derivatives for 4” Tail Chord� GOTOBUTTON _Toc387415203 � PAGEREF _Toc387415203 �45��16.4.2 Trim Angles for Varying Flight Configurations� GOTOBUTTON _Toc387415204 � PAGEREF _Toc387415204 �46��16.5 Structures� GOTOBUTTON _Toc387415205 � PAGEREF _Toc387415205 �47��16.5.1 Material Properties� GOTOBUTTON _Toc387415206 � PAGEREF _Toc387415206 �48��16.5.2 Summary of ANSYS Results� GOTOBUTTON _Toc387415207 � PAGEREF _Toc387415207 �49��16.6 Weights and Balances� GOTOBUTTON _Toc387415208 � PAGEREF _Toc387415208 �50��16.6.1 Predicted vs. Actual Weights� GOTOBUTTON _Toc387415209 � PAGEREF _Toc387415209 �51��16.6.2 Weight and Balance Summary� GOTOBUTTON _Toc387415210 � PAGEREF _Toc387415210 �52��16.7 Performance� GOTOBUTTON _Toc387415211 � PAGEREF _Toc387415211 �53��16.7.1 Parameter Values� GOTOBUTTON _Toc387415212 � PAGEREF _Toc387415212 �54��16.7.2 Performance Summary� GOTOBUTTON _Toc387415213 � PAGEREF _Toc387415213 �55��16.8 Construction� GOTOBUTTON _Toc387415214 � PAGEREF _Toc387415214 �55��16.9 Flight Testing� GOTOBUTTON _Toc387415215 � PAGEREF _Toc387415215 �56��16.9.1 Takeoff Settings for Individual Flights� GOTOBUTTON _Toc387415216 � PAGEREF _Toc387415216 �57��16.10 Modifications and Repair� GOTOBUTTON _Toc387415217 � PAGEREF _Toc387415217 �58��16.10.1 Effects of Vertical Fins� GOTOBUTTON _Toc387415218 � PAGEREF _Toc387415218 �59��16.10.2 Effect of Servo Boxes� GOTOBUTTON _Toc387415219 � PAGEREF _Toc387415219 �60��16.10.3 Aerodynamic Coefficients of Final Configuration� GOTOBUTTON _Toc387415220 � PAGEREF _Toc387415220 �61��16.10.4 Stability Derivatives 6” Tail Chord� GOTOBUTTON _Toc387415221 � PAGEREF _Toc387415221 �62��16.11 Budget Overview� GOTOBUTTON _Toc387415222 � PAGEREF _Toc387415222 �63��16.11.1 Initial and Repair Budgets� GOTOBUTTON _Toc387415223 � PAGEREF _Toc387415223 �64��17. Figures� GOTOBUTTON _Toc387415224 � PAGEREF _Toc387415224 �65��17.1 Int� GOTOBUTTON _Toc387415225 � PAGEREF _Toc387415225 �65��17.2 Design History� GOTOBUTTON _Toc387415226 � PAGEREF _Toc387415226 �66��17.2.1 Overall Configuration with 4” Tail� GOTOBUTTON _Toc387415227 � PAGEREF _Toc387415227 �67��17.2.2 Overall Configuration with 6” Tail� GOTOBUTTON _Toc387415228 � PAGEREF _Toc387415228 �68��17.3 Wind Tunnel Testing� GOTOBUTTON _Toc387415229 � PAGEREF _Toc387415229 �69��17.3.1 Wind Tunnel Testing Setup� GOTOBUTTON _Toc387415230 � PAGEREF _Toc387415230 �70��17.3.2 Aircraft Mesh with Deflectors and Resized Flaperons� GOTOBUTTON _Toc387415231 � PAGEREF _Toc387415231 �71��17.3.3 Pressure Contour with Deflector� GOTOBUTTON _Toc387415232 � PAGEREF _Toc387415232 �72��17.3.4 Pressure Contour with Ailerons� GOTOBUTTON _Toc387415233 � PAGEREF _Toc387415233 �73��17.4 Stability and Controls� GOTOBUTTON _Toc387415234 � PAGEREF _Toc387415234 �74��17.4.1 Aileron and Deflector Comparison for 0.1 rad Doublet� GOTOBUTTON _Toc387415235 � PAGEREF _Toc387415235 �75��17.4.2 Block Diagram of Yaw Damper� GOTOBUTTON _Toc387415236 � PAGEREF _Toc387415236 �76��17.4.3 Yaw Rate Comparison for Baseline and Yaw Controller� GOTOBUTTON _Toc387415237 � PAGEREF _Toc387415237 �77��17.4.4 Block Diagram of Non-linear Servo� GOTOBUTTON _Toc387415238 � PAGEREF _Toc387415238 �78��17.5 Structures� GOTOBUTTON _Toc387415239 � PAGEREF _Toc387415239 �79��17.5.1 Fuselage Internal Structure� GOTOBUTTON _Toc387415240 � PAGEREF _Toc387415240 �80��17.5.2 Finite Element Mesh of Wing� GOTOBUTTON _Toc387415241 � PAGEREF _Toc387415241 �81��17.5.3 Wing Tip Deflection at ( = 0(, 1g Loading� GOTOBUTTON _Toc387415242 � PAGEREF _Toc387415242 �82��17.5.4 Wing Tip Deflection at ( = 0(, 6g Loading� GOTOBUTTON _Toc387415243 � PAGEREF _Toc387415243 �83��17.5.5 Wing Tip Deflection at ( = 5(, 1g Loading� GOTOBUTTON _Toc387415244 � PAGEREF _Toc387415244 �84��17.5.6 Wing Tip Deflection at ( = 5(, 6g Loading� GOTOBUTTON _Toc387415245 � PAGEREF _Toc387415245 �85��17.5.7 Wing Tip Deflection with (f = 5(� GOTOBUTTON _Toc387415246 � PAGEREF _Toc387415246 �86��17.5.8 Wing Tip Deflection with (f = 30(� GOTOBUTTON _Toc387415247 � PAGEREF _Toc387415247 �87��17.5.9 (y at ( = 0(, 1g Loading� GOTOBUTTON _Toc387415248 � PAGEREF _Toc387415248 �88��17.5.10 (y at ( = 0(, 6g Loading� GOTOBUTTON _Toc387415249 � PAGEREF _Toc387415249 �89��17.5.11 (y at ( = 5(, 1g Loading� GOTOBUTTON _Toc387415250 � PAGEREF _Toc387415250 �90��17.5.12 (y at ( = 5(, 6g Loading� GOTOBUTTON _Toc387415251 � PAGEREF _Toc387415251 �91��17.5.13 (y at (f = 5(� GOTOBUTTON _Toc387415252 � PAGEREF _Toc387415252 �92��17.5.14 (y at (f = 30(� GOTOBUTTON _Toc387415253 � PAGEREF _Toc387415253 �93��17.5.15 Von Mises Stress at ( = 0(, 1g Loading� GOTOBUTTON _Toc387415254 � PAGEREF _Toc387415254 �94��17.5.16 Von Mises Stress at ( = 0(, 6g Loading� GOTOBUTTON _Toc387415255 � PAGEREF _Toc387415255 �95��17.5.17 Von Mises Stress at ( = 5(, 1g Loading� GOTOBUTTON _Toc387415256 � PAGEREF _Toc387415256 �96��17.5.18 Von Mises Stress at ( = 0(, 6g Loading� GOTOBUTTON _Toc387415257 � PAGEREF _Toc387415257 �97��17.5.19 Von Mises Stress at (f = 5(� GOTOBUTTON _Toc387415258 � PAGEREF _Toc387415258 �98��17.5.20 Von Mises Stress at at (f = 30(� GOTOBUTTON _Toc387415259 � PAGEREF _Toc387415259 �99��17.6 Weights and Balances� GOTOBUTTON _Toc387415260 � PAGEREF _Toc387415260 �100��17.6.1 Internal Component Layout� GOTOBUTTON _Toc387415261 � PAGEREF _Toc387415261 �101��17.7 � GOTOBUTTON _Toc387415262 � PAGEREF _Toc387415262 �102��17.8 Construction� GOTOBUTTON _Toc387415263 � PAGEREF _Toc387415263 �102��17.8.1 Sample Fuselage Cross-Section� GOTOBUTTON _Toc387415264 � PAGEREF _Toc387415264 �103��17.8.2 Completed Templates and Cross-Section Layout� GOTOBUTTON _Toc387415265 � PAGEREF _Toc387415265 �104��17.8.3 Templates Mounted on Blue Foam� GOTOBUTTON _Toc387415266 � PAGEREF _Toc387415266 �104��17.8.4 Joined Plug� GOTOBUTTON _Toc387415267 � PAGEREF _Toc387415267 �105��17.8.5 Top Half of Mold� GOTOBUTTON _Toc387415268 � PAGEREF _Toc387415268 �106��17.8.6 Original Single Spar Wing Cross-Section� GOTOBUTTON _Toc387415269 � PAGEREF _Toc387415269 �107��17.8.7 Wing Layout� GOTOBUTTON _Toc387415270 � PAGEREF _Toc387415270 �108��17.8.8 Wing Servo Fairings� GOTOBUTTON _Toc387415271 � PAGEREF _Toc387415271 �109��17.8.9 Wing Mounting System� GOTOBUTTON _Toc387415272 � PAGEREF _Toc387415272 �110��17.8.10 Ring-Tail Foam Core Template� GOTOBUTTON _Toc387415273 � PAGEREF _Toc387415273 �111��17.8.11 Ring-Tail Facet Cross-Section� GOTOBUTTON _Toc387415274 � PAGEREF _Toc387415274 �112��17.8.12 Ring-Tail Mounting Mechanism� GOTOBUTTON _Toc387415275 � PAGEREF _Toc387415275 �113��17.8.13 Ring-Tail Facets in Assembly Template� GOTOBUTTON _Toc387415276 � PAGEREF _Toc387415276 �114��17.8.14 Paint Scheme� GOTOBUTTON _Toc387415277 � PAGEREF _Toc387415277 �115��17.8.15 Fuselage After Coat of White Paint� GOTOBUTTON _Toc387415278 � PAGEREF _Toc387415278 �116��17.8.16 Wing After Coat of White Paint� GOTOBUTTON _Toc387415279 � PAGEREF _Toc387415279 �117��17.8.17 Completed Raven Aircraft� GOTOBUTTON _Toc387415280 � PAGEREF _Toc387415280 �118��17.9 Flight Testing� GOTOBUTTON _Toc387415281 � PAGEREF _Toc387415281 �119��17.9.1 Launch Dolly� GOTOBUTTON _Toc387415282 � PAGEREF _Toc387415282 �120��17.9.2 Aircraft Prepared for Takeoff� GOTOBUTTON _Toc387415283 � PAGEREF _Toc387415283 �121��17.9.3 First Flight Attempt Immediately After Takeoff� GOTOBUTTON _Toc387415284 � PAGEREF _Toc387415284 �122��17.9.4 Wing Damage Sustained After First Flight Attempt� GOTOBUTTON _Toc387415285 � PAGEREF _Toc387415285 �123��17.9.5 Ring-Tail Damage Sustained After First Flight Attempt� GOTOBUTTON _Toc387415286 � PAGEREF _Toc387415286 �124��17.9.6 Fuselage Damage Sustained After First Flight Attempt� GOTOBUTTON _Toc387415287 � PAGEREF _Toc387415287 �125��17.10 Modifications and Repair� GOTOBUTTON _Toc387415288 � PAGEREF _Toc387415288 �126��17.10.1 Exploded View of Wing� GOTOBUTTON _Toc387415289 � PAGEREF _Toc387415289 �127��17.10.2 Pressure Contour of Sideslip with Fins� GOTOBUTTON _Toc387415290 � PAGEREF _Toc387415290 �128��17.10.3 Aircraft Mesh with Servo Fairings� GOTOBUTTON _Toc387415291 � PAGEREF _Toc387415291 �129��17.10.4 Baseline Pressure Contour with Servo Fairings� GOTOBUTTON _Toc387415292 � PAGEREF _Toc387415292 �130��17.10.5 Baseline Pressure Contour with 6” Tail� GOTOBUTTON _Toc387415293 � PAGEREF _Toc387415293 �131��17.10.6 Ground Effect Survey� GOTOBUTTON _Toc387415294 � PAGEREF _Toc387415294 �132��17.10.7 Pitch Angle Comparison for 0.1 rad Doublet� GOTOBUTTON _Toc387415295 � PAGEREF _Toc387415295 �133��17.10.8 Pitch Rate Comparison for 0.1 rad Doublet� GOTOBUTTON _Toc387415296 � PAGEREF _Toc387415296 �134��17.10.9 Yaw Angle Comparison of Ailerons for 0.1 rad Doublet� GOTOBUTTON _Toc387415297 � PAGEREF _Toc387415297 �135��17.10.10 Yaw Angle Comparison of Deflectors for 0.1 rad Doublet� GOTOBUTTON _Toc387415298 � PAGEREF _Toc387415298 �136��17.10.11 Roll Angle Comparison of Ailerons for 0.1 rad Doublet� GOTOBUTTON _Toc387415299 � PAGEREF _Toc387415299 �137��17.10.12 Roll Angle Comparison of Deflectors for 0.1 rad Doublet� GOTOBUTTON _Toc387415300 � PAGEREF _Toc387415300 �138��17.10.13 Wing Pivot Mechanism� GOTOBUTTON _Toc387415301 � PAGEREF _Toc387415301 �139��18. Index� GOTOBUTTON _Toc387415302 � PAGEREF _Toc387415302 �140����Introduction

The 1996-97 Senior Design Project in aerospace engineering at North Carolina State University was the design, construction, and flight testing� XE "Flight test" � of a remotely piloted vehicle (RPV). The ultimate goal of this project is to serve as a prototype for a tube-launched reconnaissance aircraft. This document is a report of the construction and flight testing of the Raven, one possible solution to this unique problem. All written work and drawings referencing the design of the Raven can be found in Volume I (Fall semester, 1996).

It is important to note the set of constraints placed upon the Raven aircraft at the outset of the design process, since these constraints significantly impacted the construction of this aircraft. The constraints were as follows:

Ring-tail empennage, pivoting at least in pitch� XE "Pitch" �Maximum stowed radial dimension of 7”Maximum 60” wing� XE "Wing" � spanMinimal weight� XE "Weight" �Design loads� XE "Design loads" �� XE "Load" � of +4 and -2Safety factor� XE "Safety factor" � of 1.510% power-off static margin� XE "Static margin" �OS Max 61SFABC-P engine� XE "Engine" �/tractor propeller� XE "Propeller" �Swing-wing� XE "Wing" � upgrade capability (for tube-launching)Dolly� XE "Dolly" � launch with skid� XE "Skids" � landing� XE "Landing" �

Total Quality Management (TQM) was a primary goal in this design effort. Maximizing productivity and creativity through effective communication and well-planned progress was a top priority. Concurrent engineering techniques were used to lay down the framework for this progress as outlined in Volume I.

An overall configuration� XE "Configuration" � was developed early on, with the flexibility to allow changes when necessary. During flight testing� XE "Flight test" �, this flexibility in design would prove very important. This configuration was based on a commitment to meet or exceed all of the design criteria.�Design History

At the end of the fall semester, the overall configuration� XE "Configuration" � of the aircraft, described in detail in Volume 1 of this work, included a number of distinguishing characteristics. The driving force in the design, the ring-tail, was four inches in diameter and comprised of 12 individual facets. The wing� XE "Wing" � was an SD7062 airfoil with a 6.4 inch chord� XE "Chord" � and five foot wing span. The control surfaces� XE "Control surfaces" � on the wing consisted of flaperons for roll� XE "Roll" � control� XE "Flaperons" � and drag� XE "Drag" � rudders� XE "Drag rudders" � for yaw� XE "Yaw" � control (see Volume I for a more complete description). It took many iterations to reach this configuration in the fall, so it was not surprising when other changes took place during the spring semester. The aircraft configuration prior to these changes is shown in figure 17.2.1. Some of the changes made were a result of new analysis conducted in the spring, others were a result of efforts to ease construction, and a few were the results of flight testing� XE "Flight test" �. Yet, as in the fall semester, the driving factor throughout everything was the ring-tail. The final configuration of the spring semester is shown in figure 17.2.2.

Wing� XE "Wing" �

At the end of the fall semester, PMARC� XE "PMARC" � analysis revealed that the flaperons� XE "Flaperons" � were producing a strong side force that would make it difficult to control the aircraft. Further investigation of this phenomenon in the wind tunnel revealed that it might be possible to use this yaw� XE "Yaw" � force to our advantage through the use of control surfaces� XE "Control surfaces" � called deflectors� XE "Deflectors" �. Stability� XE "Stability" � and controls analysis in the fall had already determined that the drag� XE "Drag" � rudders� XE "Drag rudders" � would not provide sufficient yaw control for the aircraft. The deflectors were small, inboard flaps� XE "Flaps" � that deflected in the up direction only. For example, an upwards deflection� XE "Deflection" � of the right deflector would produce a suction on the right side of the tail, causing the plane to yaw to the left. The flaperons were subsequently moved farther outboard and sized based on PMARC analysis that suggested they would still provide adequate roll� XE "Roll" � control without interfering with yaw control.

The wing� XE "Wing" � also underwent a structural evolution in the spring semester. Originally, the wing spar� XE "Spars" � was to be made out of wood. However, little analysis was done on this configuration� XE "Configuration" � in the fall due to the unavailability of ANSYS, the finite element analysis software package. In the course of the spring semester, analysis proved the wing loading� XE "Wing loading" � to be too high for a wood spar at the upper limit load� XE "Load" � of +6g. Therefore, a strong, yet lightweight alternative was needed. This alternative was a C-channel carbon� XE "Carbon" � fiber spar placed at the quarter chord� XE "Chord" �. Because of carbon’s high strength, this eliminated the need for a trailing edge spar, but it did not make room for the wing servos� XE "Servos" � to fit inside the wing. Even near the quarter chord point, the wing was not thick enough to allow for the servos to fit inside it. Therefore, servo fairings� XE "Fairing" � were also constructed to help reduce the drag� XE "Drag" � on the servos. This was the set configuration when construction began and during the first flight attempt. During the subsequent crash, the wing was completely destroyed, and one theory behind the crash was that the servo fairings had somehow interfered with the flow through the tail. Therefore they were removed. To do this, the spar was moved forward to the leading edge, and a second, smaller carbon spar was added to the trailing edge. This, along with a slight increase in wing thickness, allowed the servos to fit completely inside the wing. This was the wing configuration wing for all subsequent flights.

Tail

Although the tail was the one thing that remained a constant throughout the construction process, this was not the case during flight testing� XE "Flight test" �. After the wing� XE "Wing" � servo� XE "Servos" � fairings� XE "Fairing" � were removed, it was assumed that the aircraft would fly as predicted. This was not the case, and the aircraft crashed during the second flight attempt. Although all analysis predicted the aircraft should fly, it was believed that the four inch ring-tail did not provide enough control power in a real flight situation. The solution to this problem was to use a larger tail, specifically a perfectly cylindrical tail with a six inch diameter. Following construction of this tail, a third flight attempt was made. This was the first successful flight of the aircraft, and the tail remained unchanged on subsequent flights.

Skids� XE "Skids" �

During the third and fourth flights, the original main skid� XE "Skids" � design proved to be unable to sustain the landing� XE "Landing" � loads� XE "Load" �, which were estimated to be much higher than the design load� XE "Design loads" � of 6g’s. Therefore, a new approach was utilized to make the skids more durable. The basic geometry and materials were left unchanged, but the new attach mechanism incorporated machine nuts and washers for quick repairs. The wing� XE "Wing" � tip� XE "Wing tip" � skids and the tail skids withstood all flights and were therefore left unchanged.�Wind Tunnel Testing

The first aerodynamic� XE "Aerodynamics" � analysis in the spring semester involved an investigation of the adverse aileron� XE "Ailerons" � yaw� XE "Yaw" �� XE "Configuration" �, which is described in detail in volume 1. Wind tunnel testing was conducted using a full-size fuselage� XE "Fuselage" � and a wing� XE "Wing" � that spanned the test section of the tunnel as shown in figure 17.3.1. By using a tufted wand to trace the flow trailing from the control surfaces� XE "Control surfaces" �, the wind tunnel test showed that the adverse yaw� XE "Adverse yaw" � originated from vortices� XE "Vortex" � on the inboard edges of the ailerons. Thus, by moving the ailerons outboard and using small inboard flaps, called deflectors� XE "Deflectors" �,� XE "Flaps" � that deflected opposite the ailerons, the effect was overcome. In addition, because moving the ailerons outboard placed them further from the tail, their influence over the tail decreased. The deflectors were placed directly upstream of the tail, so they had a much stronger influence than the ailerons. Thus, when deflected opposite the ailerons, the deflectors outweighed the adverse aileron yaw and produced a positive aileron yaw while being far enough inboard and small enough to avoid significantly compromising the roll� XE "Roll" � control of the ailerons. In fact, the side force on the tail appeared to be directly proportional to the deflector angle at zero sideslip. Furthermore, the deflectors worked best when only one at a time was deflected upward. Wind tunnel testing also revealed that no spacing between the deflectors and flaperons� XE "Flaperons" � is necessary to keep the flaperon vortices from interfering with the deflector vortex, and that the deflector vortex is speed-sensitive, moving downward with decreasing speed. The deflectors proved powerful enough to be feasible for yaw control, a replacement for the drag� XE "Drag" � rudders� XE "Drag rudders" � that had already been deemed inadequate. Figure 17.3.2 shows the revised control surface layout with deflectors and resized flaperons. Figure 17.3.3 gives the deflector pressure contour and figure 17.3.4 displays the new aileron pressure contour. The deflector contour shows suction inside the tail that is absent in the aileron contour, but would produce a positive aileron yaw. The table of all of the aerodynamic coefficients� XE "Aerodynamic coefficients" � described in the design modifications section� XE "section" � gives the data from these cases. Note that the deflectors can compensate for the side force and yawing moment of the flaperons.

Next, PMARC� XE "PMARC" � analyses for pitch� XE "Pitch" � and yaw� XE "Yaw" � rate had to be redone. Last semester's pitch and yaw rate cases had used the nose of the aircraft as the rotation point, due to the setup of PMARC's coordinate system. The rotation point was moved to the center of gravity, the natural point of rotation for a solid body, and these cases were run again. The new pitch and yaw rate coefficients are given in table 16.3.1. The other coefficients remain unchanged from those given in Volume 1.

�Stability� XE "Stability" � and Controls

Drag� XE "Drag" � Rudder to Deflectors� XE "Deflectors" �

As discussed in Section 3, the ineffective drag� XE "Drag" � rudders� XE "Drag rudders" � were replaced with deflectors� XE "Deflectors" �. The deflectors were modeled in PMARC� XE "PMARC" � to obtain the coefficients needed for dynamic simulations. However, the data from the wind tunnel tests provided the side force coefficient for the deflectors. The coefficients are listed in Table 16.4.1. After running a simulation the deflectors proved to have more control power than the drag rudders, but not more than the ailerons� XE "Ailerons" �. This is shown in Figure 17.4.1 in which the response of the deflectors is compared to that the ailerons. This was confirmed during flight testing� XE "Flight test" � when the deflector was used to yaw� XE "Yaw" � the aircraft. It had little response to the deflector in yaw, but a roll� XE "Roll" � response was clearly seen.

Controller� XE "Controller" �

Another requirement for the project was to design a yaw� XE "Yaw" � damper to control the aircraft. This was done in conjunction with the MAE 525 class. The block diagram of this can be seen in Figure 17.4.2. A group of three people worked on the controller� XE "Controller" � for this particular aircraft. While each person worked on identical controllers, the amount of damping on the Dutch Roll� XE "Roll" � for the yaw rate differed for each. A lightly, moderately, and heavily damped controller were all utilized. Figure 17.4.3 shows responses for the aircraft without the controller and with a lightly damped controller.

In order to design the controller� XE "Controller" �, a continuous time system was developed which included aircraft dynamics, servo� XE "Servos" �, and washout circuit. The gain value and time constant were found by plotting several root loci for varying time constants. Once the time constant was found, the gain could be found for the chosen amount of damping of the Dutch Roll� XE "Roll" �. For example, the gain for the lightly damped controller was 0.2803 and the time constant was 3.0 seconds. The system was then converted to a discrete time system to determine the discrete gain and discrete time constant. These were compared to the continuous time values for validity of the system.

The controller� XE "Controller" � was designed using a continuous time non-linear� XE "Linear" � servo� XE "Servos" � and aircraft dynamics along with a discrete washout circuit, rate limiter, and time sampler (time sampled every 0.043 seconds). The non-linear servo is shown in Figure 17.4.4. A robustness check was performed on each variable of the system. It was found that velocity� XE "Velocity" �, Cn(, Cnr, and Clp were the most sensitive to changes. Noise was also added into the controller� XE "Controller" � to resemble outside disturbances. After this was completed, the simulation was checked to make sure the controller behaved as expected. Note that responses were checked throughout the design of the controller. Although the controller was designed, it was never used due to the down-time of the aircraft and the fact that the aircraft has not had a fully instrumented flight to date.

Trim� XE "Trim" �

The trim angle� XE "Trim" �s for takeoff and level flight were calculated using a self-written program found in the Appendix. The equations used were for linear� XE "Linear" � lift and moment curves and included all control surface deflections� XE "Deflection" � and incidences. The settings for different flight configurations� XE "Configuration" � are shown in Table 16.4.2.

�Structures

Structures Review

Ideally, all structural analysis should have been completed during the fall semester’s design phase. However, because the software license for ANSYS expired on the EOS computing system, that was not possible. ANSYS did not become available at the very end of the fall semester through the North Carolina Supercomputing Center (NCSC), and therefore, structural analysis was not conducted until the spring semester. In addition, because it was so late in the year when ANSYS became available, structural analysis was only conducted on the wing� XE "Wing" �.

The overall structural design of the aircraft is described in detail in Volume I. However, some changes were made in the beginning of the spring semester to facilitate construction. Originally, the fuselage� XE "Fuselage" � skin� XE "Skin" � was to be made of five layers of bi-directional Kevlar� XE "Kevlar" �. This was reduced to only three layers to decrease weight� XE "Weight" �. Additionally, the original stringers� XE "Stringers" � in the aircraft were to be made of strips of poplar� XE "Poplar" � wood. Instead, the stringers were made of uni-directional carbon� XE "Carbon" � fiber because it not only weighs less, but is also much stronger. The entire internal layout of the fuselage, including the stringers, can be seen in figure 17.5.1.

The only major structural change in the spring semester concerned the wing� XE "Wing" �. Originally, the wing had been designed with a single wooden spar� XE "Spars" � located at the quarter chord� XE "Chord" �. But preliminary analysis in the spring semester revealed that this spar would be unable to withstand applied loads� XE "Load" � at +6g, the upper load limit. Therefore, a stronger, yet lightweight alternative was needed. This alternative came in the form of a C-channel spar made from two layers of bi-directional carbon� XE "Carbon" � fiber laid at 90( located again at the quarter chord. This was the configuration� XE "Configuration" � of the wing for the first flight. However, as a result of a modification made during flight testing� XE "Flight test" �, the spar had to be moved forward to the leading edge of the wing (see section� XE "section" � 10). Because the spar was then so far forward of the quarter chord point, a trailing edge spar was added to increase the torsional stiffness of the wing. While the forward spar was still made of two layers of bi-directional carbon, the trailing edge spar was made of only one layer. The wing analysis discussed in this section was performed on the two-spar wing.

Finite Element Analysis

The two primary tools used in the analysis were Unigraphics and ANSYS, both described in Volume 1. Because the wing� XE "Wing" � is symmetric about the aircraft centerline, it was only necessary to model half the span. The wing mesh� XE "Meshes" �, shown in figure 17.5.2, is basically a balsa� XE "Balsa" � wood shell with two carbon� XE "Carbon" � fiber spars� XE "Spars" � running through it at the leading and trailing edges (see table 16.5.1 for material properties� XE "Material properties" �). The 2480 elements in this mesh are spaced closer together at the wing root because in these areas the stresses� XE "Stress" � change more rapidly, and making the elements smaller yields a more accurate analysis. In figure 17.5.2, as well as several other ANSYS figures, the wing root is at the bottom of the page, wing tip� XE "Wing tip" � at the top, and the leading edge is on the left.

The wing� XE "Wing" � was analyzed at 50 mph under both 1g and 6g loads� XE "Load" � for ( = 0( and 5(, as well as with 5( and 30( flap� XE "Flaps" � deflections� XE "Deflection" �. The 6g load cases were found by simply multiplying the 1g load cases by six, and the 30( flap deflection load was found by linearly interpolating to 30( using the 0( and 5( flap deflection cases. To simulate the restrictions on the wing in flight, the nodes at the root of the wing were constrained in all six degrees of freedom, as were nodes that coincided with the location of the attachment for the wing mounting� XE "Mounting" � system� XE "Wing mounting system" �.

In the analysis, the focus was on both the wing� XE "Wing" � tip� XE "Wing tip" � deflection� XE "Deflection" �, (y (stress� XE "Stress" � in the y plane), and the Von Mises stress, the average of the stresses in all three planes found from the equation (vm = ((x2 + (y2 + (z2)½. Strain� XE "Strain" � was not closely examined due to the fact that ANSYS has some problems with Poisson’s Ratio for balsa� XE "Balsa" � wood, and there was no way to be certain if the strain values shown were correct.

The overall results of the finite element analysis are shown in table 16.5.2. At zero angle of attack, the wing� XE "Wing" � tip� XE "Wing tip" � deflections� XE "Deflection" � are 0.27” under 1g loading and 1.613” under 6g loading. Looking also at the 5( angle of attack case it is clear that these numbers do not increase significantly, nor do they increase when the flaps� XE "Flaps" � are deflected 5(. This is also clearly shown in figures 17.5.3 - 17.5.7. However, looking at figure 17.5.8, a 30( flap deflection results in a negative wing tip deflection of -0.381”. In all other cases, the wing flexed up, but in this case the wing flexed down. This is due to the increased lift near the root of the wing caused by the flap deflection. This bows the wing at the root, forcing the root up and the tip down.

The bending of the wing� XE "Wing" � causes compressive and tensile stresses� XE "Stress" � in the y plane, denoted as (y. It is these stresses that are the highest and will cause the wing to fail if the loading becomes too high. When the wingtip is deflected in the positive direction, the upper surface of the wing is in compression� XE "Compression" � while the lower surface is in tension� XE "Tension" �. Compressive stresses are denoted as negative, so it is apparent from table 16.5.2 that the maximum (y in most all of the cases results from compression. The areas of highest stress occurred on the backside of the leading edge spar� XE "Spars" � at the wing root. This is the area on the spar that is connected in with the wing mount� XE "Mounting" � system, The stress contour can be seen in figure 17.5.9, showing (y for the zero angle of attack case under 1g loading. In this figure, the wing root cross-section� XE "section" � is seen with the backside of the leading edge spar. Here, the maximum stress is -12732 psi and is shown as the black area on the graph. Under 6g loading at zero angle of attack, (y increases to -76392 psi, as seen in figure 17.5.10. Figures 17.5.11 - 17.5.14 show the results from the other load� XE "Load" � cases, however, the 6g loading at zero angle of attack shows the highest stress levels, but it is not so high as to cause concern, since the ultimate compressive and tensile stresses of carbon� XE "Carbon" � fiber are both 180,000 psi. One might also notice that like the wing tip� XE "Wing tip" � deflection� XE "Deflection" � results, the numbers for the 30( flap� XE "Flaps" � deflection (figure 17.5.14) differ in sign from the other cases. Once again, the high lift near the root of the wing causes the wing to bow upwards at the root, thereby placing upper surface of the wing at the root in tension rather than compression. This, of course, results in a positive value for (y.

The stresses� XE "Stress" � in the x and z planes are an order of magnitude less than those in the y plane, so it is not necessary to examine them individually. However, it is worth looking at the overall average of the stresses, the Von Mises stress. Because the Von Mises stress is an average stress, it’s values are always positive. As seen in table 16.5.2, the Von Mises stress is not much greater than (y, again indicating that (x and (z are much smaller than (y. The values also are less than half the ultimate stress of carbon fiber, so it is safe to say the wing� XE "Wing" � can handle the applied loads� XE "Load" �. The ANSYS results for Von Mises stress are shown in more detail in figures 17.5.15 - 17.5.20.

As mentioned previously, the second, trailing edge spar� XE "Spars" � was added to the wing� XE "Wing" � to increase it’s torsional stiffness. It is then important to note that throughout this analysis, the rotation of the wing was negligible, even in the 30( flap� XE "Flaps" � deflection� XE "Deflection" � case.

Conclusions

Based on the analysis shown here, the two-spar� XE "Spars" � wing� XE "Wing" � is sufficient to withstand the applied aerodynamic� XE "Aerodynamics" � loads� XE "Load" � with the current wing mount� XE "Mounting" � system. The highest stress� XE "Stress" � concentrations are seen on the leading edge spar where the spar is connected to the wing mount, but these stresses are still less than half the ultimate stress of carbon� XE "Carbon" �. This analysis is supported through some qualitative flight test� XE "Flight test" � information. During flight testing, there have been several landings� XE "Landing" � that have been estimated to be over 6g, and to date the wing has suffered no damage. In addition, this analysis excludes the blue foam� XE "Foam" � in the wing as well as the fiberglass� XE "Fiberglass" � on the outside of the wing, which makes the analysis conservative from the start. �Weights� XE "Weight" � and Balances

At the end of last semester the estimated dry aircraft weight� XE "Weight" � was 12.05 lbs. The actual dry weight of the aircraft at the time of the first flight was 9.9 lbs. The reason for this difference was the over-estimation of the weight of several unknowns, primarily resin� XE "Resin" �, paint, and wiring. Another reason was the loss of almost 1.2 lbs from the lack of two digital cards� XE "Digital cards" � and the flight data recorder (FDR� XE "FDR" �). A comparison of the predicted weight and actual weight is available in Table 16.6.1

In addition, the fact that the digital cards� XE "Digital cards" � and FDR� XE "FDR" � were not in the aircraft at first flight required other components to be moved to achieve the required static margin� XE "Static margin" � of 10% power off. The battery packs were moved aft in order to facilitate this. The smaller battery pack was moved to directly behind the SAS� XE "Stability Augmentation System (SAS)" � (Stability� XE "Stability" � Augmentation System), and the larger battery pack was moved behind the opto-isolater card. A 0.1 lb ballast was added in front of the tail skid� XE "Skids" � before the first flight to balance the plane longitudinally. The component layout used to reach the required static margin, minus the ballast weight� XE "Weight" �, is shown in figure 17.6.1. No weight was required for lateral balance.

The aircraft dry weight� XE "Weight" � was 9.5 lbs after reconstruction of the wing� XE "Wing" �. The weight reduction was attributed to the loss of fiberglass� XE "Fiberglass" � and resin� XE "Resin" �. A 0.2 lb ballast was again added in the same location before the second flight to balance longitudinally. The increase in ballast weight is attributed to the additional resin and fiberglass on the front of the fuselage� XE "Fuselage" � due to repair work. Once again, no ballast weight was needed laterally.

With the introduction of the new tail and more resin� XE "Resin" � and fiberglass� XE "Fiberglass" � to the front of the fuselage� XE "Fuselage" � the aircraft dry weight� XE "Weight" � rose to 10.1 lbs after the second flight attempt. In balancing the aircraft for flight, the small battery pack was moved to the front of the aircraft, so no ballast was needed. The same configuration� XE "Configuration" � was used in the fourth flight. A quick summary of the aircraft weight at each flight is available in Table 16.6.2.

The moments of inertia� XE "Moment of inertia" � were estimated in the fall semester using Unigraphics, and these estimations are available in Volume 1. The moments of inertia were not measured in the spring because the aircraft was never fully instrumented, and the estimates were for the fully instrumented aircraft.

�Performance� XE "Performance" �

Analysis this semester involved recalculating the aircraft performance� XE "Performance" � characteristics (takeoff� XE "Takeoff" �, cruise, and landing� XE "Landing" �) with the new weight� XE "Weight" � of 12 lbs, because the original calculations assumed an estimated weight of 15 lbs. The 20% reduction in weight was significant enough to warrant recalculation, especially at the critical takeoff. In addition, a number of new flap� XE "Flaps" � calculations were performed. The actual control surface deflections� XE "Deflection" � for various flight conditions are presented in Section 4.

Takeoff� XE "Takeoff" � calculations began by re-running the Bungee� XE "Bungee" � program to determine the bungee pullback distance with the updated weight� XE "Weight" �. To reach the takeoff velocity� XE "Velocity" � of 70.7 ft/s (1.2*Vstall), the bungees must be pulled back 23.2 feet. At this distance, the plane will experience a maximum acceleration of 4.2 g’s. However, the Bungee program does not account for friction or wear of the bungees over time. Therefore, based on past experience, a pullback distance of 42 feet was used. A takeoff lift to weight ratio of 1.2 together with the 70.7 ft/s takeoff velocity determines a takeoff lift coefficient, CL of 0.9085. Using the parabolic drag� XE "Drag" � model� XE "Drag model" � (Table 16.7.1), this in turn determines a takeoff drag coefficient, CD, of 0.06625.

The 15 lb plane was originally designed to fly level (( = 0() at cruise (65 mph) without flaps� XE "Flaps" �. In order to achieve this condition, the wing� XE "Wing" � is mounted on the fuselage� XE "Fuselage" � at a 2( incidence to provide the required CL of 0.5211. This generates a CD of 0.0419. Since CL is a function of angle of attack, the 12 lb aircraft would still generate a CL of 0.5211 when flying level at cruise. However, the actual lift generated would no longer balance the weight� XE "Weight" �. In fact, 20% extra lift would be generated causing the plane to climb. For lift to still balance the weight, the 12 lb aircraft will require a CL of 0.4169 with a CD of 0.03763. But to fly level, the flap deflections� XE "Deflection" � will no longer be zero.

Finally, the original landing� XE "Landing" � calculations were performed by a FORTRAN program which calculated gliding performance� XE "Performance" � at various velocities� XE "Velocity" � using the non-dimensional CL and CD alone. Therefore, the aerodynamic� XE "Aerodynamics" � coefficients� XE "Aerodynamic coefficients" � at the actual aircraft weight� XE "Weight" � remained the same. At approach, (1.3*Vstall), CL is 0.8044 and CD is 0.05841. And at touchdown (1.15*Vstall), CL is 1.027 and CD is 0.07632. The parameters used in the performance calculations are presented in table 16.7.1, and takeoff� XE "Takeoff" �, cruise, and landing performance are summarized in table 16.7.2.

�� XE "Performance" �Construction

Fuselage� XE "Fuselage" �

Plug� XE "Plug" �

The first step in the fuselage� XE "Fuselage" � construction was the formation of the plug� XE "Plug" �. The shape of the plug would determine the shape of the RavenÕs fuselage, and so throughout the construction of the plug, care was taken to ensure accuracy. The first step in the plug construction was to fabricate a set of templates which would later be used to define the shape of the plug. To begin, cross-sections were generated in Unigraphics. The distribution of cross-sections along the length of the fuselage was a balance between ease of construction and accuracy of the shape. The cross-sections were placed two inches apart along relatively flat areas which could be easily shaped and one inch apart along quickly sloping areas to better maintain the correct curvature. Figure 17.8.1 shows a sample cross-section� XE "section" �. Hard copies of the cross-sections were then cut in half along the horizontal center line. Only half of the template� XE "Template" � was required, because except for the wing� XE "Wing" � mount� XE "Mounting" � on the top half, the RavenÕs top and bottom halves were mirror images of each other. To take advantage of this, the top and bottom halves of the plug were constructed separately from identical sets of Formica� XE "Formica" � templates. Next, the cut out cross section halves were attached to sheets of Formica using 3M-77 spray adhesive, from here on referred to as 77. Formica was chosen because it is difficult to sand and would therefore maintain its shape during the upcoming foam� XE "Foam" � sanding process. Both top and bottom templates were cut out simultaneously to ensure they would be identical. The templates were then roughly cut to shape and finally sanded down to the paper cross section. When the desired shape was reached, notches were cut into the edge to mark the vertical center line. This was an attempt to define another reference line on the plug which would then be transferred to the mold� XE "Mold" � and ultimately to the finished fuselage skin� XE "Skin" �. The finished templates and the cross-section layout are shown in Figure 17.8.2.

The next step in the plug� XE "Plug" � construction was to fabricate the blue foam� XE "Foam" � blocks that would make up the actual plug. First, two pairs of aluminum rails were cut to necessary heights, 31/32Ó and 131/32Ó, to be used as a guide during the hot-wiring process. The Unigraphics distribution placed cross section� XE "section" � 1Ó and 2Ó apart, but to account for the 1/32Ó thickness of the Formica� XE "Formica" � templates, the foam blocks had to be hot wired� XE "Hot wire" � to 31/32Ó and 131/32Ó. After hot wiring, the templates were bonded to the foam with 77. The last step in constructing the foam blocks was to roughly cut the foam to shape to decrease the amount of sanding later. This is illustrated in Figure 17.8.3.

With the foam� XE "Foam" � blocks finished, assembly of the plug� XE "Plug" � began. A door was used as a table because a flat surface was required to build the two halves of the plug. A full size Unigraphics printout of the cross-section� XE "section" � layout was taped to the table. This was used to align the foam blocks. A second print out was used to cut the correct fuselage� XE "Fuselage" � shape into wood Òshape testersÓ which would be used to check the shape of the plug during the sanding process. Before gluing the foam blocks together, they were checked with the print out and sanded to the correct thickness when necessary. The blocks were individually aligned with the print out and then glued together. After allowing the glue to dry thoroughly, the foam was sanded to shape. Long strokes with sanding blocks were used to avoid sanding down too far. One unforeseen problem was that the blue foam in the two plug halves shrank overnight due to weather changes. Foam had to be replaced in the half that shrank more to realign the templates in the two halves. Immediately after realigning the templates, the two halves were joined with wood glue to avoid further differential shrinking problems. Figure 17.8.4 shows the joined plug. Spackle was applied and sanded down to fill any small gaps or dings and to raise low areas in the plug. Then, the entire plug was spackled and sanded smooth to seal it. A single layer of six ounce fiberglass� XE "Fiberglass" � was wrapped around the plug to aid in smoothing the surface. After curing, the plug was sanded smooth. With the fiberglass to protect the blue foam, Bondo� XE "Bondo" � was applied to the plug and sanded down. Then the plug was primed. Several iterations of Bondo, sand, prime, and sand were performed to perfect the plugÕs shape. By this time, the original notches cut into the templates to mark the centerline of the plug had been covered with spackle, glass, and Bondo so they could no longer be used to generate a reference line on the plug. A new method to mark reference lines directly in the mold� XE "Mold" � was found later.

After completing the plug� XE "Plug" �, mold� XE "Mold" � fabrication began. First, a cradle to hold the plug was constructed. The shape of bottom half of the plugÕs constant area section� XE "section" � was hot wired� XE "Hot wire" � out of the top of a length of blue foam� XE "Foam" � to support the plug during mold construction. The foam support was then glued to a wood base upon which the rest of the mold could be built. The height of the centerline in the front and back of the plug were measured to ensure the wedge was holding the back of the plug at the correct height. With the plug secured, the boom was attached. A 2Ó diameter tube was used as the plug boom and was attached by means of a wooden dowel� XE "Dowel" �. A separate foam block was made to support the boom. Next, a Plexiglas parting plane was roughly cut to the shape of the plug. More foam blocks were used to support the two halves of the parting plane at the correct height, along the horizontal centerline of the plug. Final shaping of the parting plane to match the curvature of the plug was done with a Dremmel tool. The two halves of the parting plane were then secured in place atop the foam supports with Bondo� XE "Bondo" �. To complete the cradle, the small gap between the plug and the parting plane was filled with body filler and sanded down.

Once the cradle was completed, the entire assembly was sprayed with PVA so the mold� XE "Mold" � would release easily. Then the actual mold was ready to be laid up. First, two layers of nine ounce fiberglass� XE "Fiberglass" � were laid up over the bottom half of the plug� XE "Plug" � and the parting plane. Then, a heavier layer of 18 ounce fiberglass was laid up over them to make the mold more rigid. After curing overnight, wooden legs were attached to the mold with body filler. Then the cradle was turned upside down to sit on these new legs and reveal the top half of the plug. The Plexiglas was removed and the bottom half of the mold became the new parting plane for the top half. The same process was repeated for the top half. After curing, the two halves were separated. Finally, wooden legs were added to the top half of the mold and Plexiglas walls to lay up on were added to the front and back of both halves yielding two separate, free standing mold halves to lay up the actual fuselage� XE "Fuselage" � skin� XE "Skin" � in. The finished top half of the mold is presented in Figure 17.8.5.

Before laying up the fuselage� XE "Fuselage" � skin� XE "Skin" �, reference lines had to be drawn in the mold� XE "Mold" � to mark the location of the stringers� XE "Stringers" � and hatches� XE "Hatches" �. References lines parallel to the fuselageÕs constant area section� XE "section" � were drawn on the ledges of the mold. A tool was constructed of foam� XE "Foam" � and Formica� XE "Formica" � to generate reference lines at 45( from vertical In addition to reference lines, the mold had to be prepared in advance for joining the two halves of the skin. In order to avoid working around the bulkheads� XE "Bulkhead" �, it was decided to join the two halves of the fuselage skin on the outside with strips of fiberglass� XE "Fiberglass" �. In order to create the channel in the fuselage skin to hold these strips, five layers of masking tape matching the thickness of two layers of two ounce fiberglass were applied along the inside surface of the mold. Then the mold was waxed and sprayed with PVA.

With the mold� XE "Mold" � prepared, the lay up of the bottom half of the fuselage� XE "Fuselage" � began. Three layers of Kevlar� XE "Kevlar" � which had been pre-cut to fit in the mold were laid up at 45(. The excess Kevlar which extended outside the mold was simply folded over onto the ledges of the mold. However, the Kevlar could not turn such a sharp corner and bubbles formed which made it difficult to get the skin� XE "Skin" � to lay flat in the mold. Another problem arose during the vacuum bagging. The vacuum bag was attached to the ledge and vacuum bag only the inside surface of the mold. But with the ledge covered with spilled epoxy� XE "Epoxy" � it was difficult to get a good seal for the vacuum. Because of this, numerous voids where bubbles had formed were found the next day. This skin was unusable, and a second lay up was required. This time, foam� XE "Foam" � walls wrapped in packing tape were erected around the mold to support the excess Kevlar extending up from the mold. Another difference was that the carbon� XE "Carbon" � fiber stringers� XE "Stringers" � were laid up between layers of Kevlar to simplify construction. And finally, to solve the vacuum problem, the entire mold was vacuum bagged. This resulted in a fuselage skin of much higher quality than the first attempt. The last step in the bottom fuselage halfÕs construction was to use the Dremmel cutting tool to remove the excess Kevlar extending up from the mold.

Because all the hatches� XE "Hatches" � are on the upper surface of the aircraft, they had to be constructed first so that the fuselage� XE "Fuselage" � skin� XE "Skin" � could be laid up on top of them. The front and back positions of the hatches were marked on the mold� XE "Mold" � ledges. The sides of all hatches were at 45( from vertical. Once again, the foam� XE "Foam" �-Formica� XE "Formica" � tool was used to mark lines inside the mold at 45(. Blue foam frames were used to lay up the frames in. In order to transfer the 3D shape of the hatches from inside the mold onto the flat pieces of foam, wax paper was laid inside the mold and the shape of the hatches was traced onto it. Wax paper was utilized because being transparent, it was simple to trace the hatch shape already drawn in the mold onto it. Then this shape was traced onto 0.5Ó thick foam, and the shape was cut out with the band saw. The inside walls of the frame were bevel cut at 45( so that the hatchesÕ ledges would also be laid up at this angle. This presented a problem later. The foam frames were then wrapped with packing tape, waxed, pressed into the waxed and PVAÕd mold, and glued in place. To begin the lay up, carbon� XE "Carbon" � strips were pressed into the inside corners of the hatches which were then filled with a mixture of epoxy� XE "Epoxy" � and microballoons. When completed, this would form the sharp edges of the hatches. This was necessary because the fiberglass� XE "Fiberglass" � would not turn the sharp corners. Two layers of six ounce glass at 45( were then used to lay up the hatches. Two carbon strips were laid up between layers of fiberglass to increase stiffness of the hatches. However, the 45( bevel of the foam frame added to the curvature of the mold created an extremely tight corner which could neither be filled with the epoxy and microballoons� XE "Microballoons" � nor the fiberglass. As a result, the finished hatches had very ragged edges which made them unusable. A second attempt using non-beveled foam frames yielded much better results. Another change made during the second hatch lay up was to move the sides of the hatches out to 50( from 45(. This was done to increase the size of the hatches for increased access to the finished fuselage internal systems� XE "Internal systems" �. To finish the hatches, their position in the mold was marked, and they were removed to round their corners and fill in any voids not filled with epoxy and microballoons during the initial lay up. The hatches were then replaced in the mold to prepare for the lay up of the upper half of the fuselage skin.

The top half of the fuselage� XE "Fuselage" � was laid up identically to the bottom half except holes had to be pre-cut to allow the Kevlar� XE "Kevlar" � to fit over the hatches� XE "Hatches" �. The holes were cut big so as to leave a gap between the Kevlar and the hatches. This gap was filled with eight ounce glass which laid up against the hatches better than Kevlar could. After curing, the hatch flanges were cut down to approximately one quarter inch.

The next step in the fuselage� XE "Fuselage" � construction, bulkhead� XE "Bulkhead" � fabrication, actually began while the skin� XE "Skin" � was being laid up. Like the fuselage templates, Unigraphics printouts were bonded to poplar� XE "Poplar" � plywood� XE "Plywood" � which was then roughly cut to shape. And like the hatches� XE "Hatches" �, the bulkheads were located in the fuselage by measuring from the nose and marking the ledge of the mold� XE "Mold" �. The bulkheads were then sanded to fit inside the fuselage one half at a time. To ensure that the bulkheads were perfectly vertical, a T-square supported by a ruler laying across the mold was butted up against them. After all of the sanding was completed, the skid� XE "Skids" � and wing� XE "Wing" � mount� XE "Mounting" � bulkheads were glassed with a single layer of six ounce fiberglass� XE "Fiberglass" � and vacuum bagged. Also during this time, all other internal structures such as the SAS� XE "Stability Augmentation System (SAS)" � mount and skid hardpoints were installed. The final installation involved bonding the bulkheads first into the bottom half of the fuselage using epoxy� XE "Epoxy" � and flocking� XE "Flocking" �, and then after drying, the top half of the fuselage was placed on top of the bulkheads and bonded down. For added strength, the bulkheads were filleted all around with epoxy and flocking. Both sides of the firewall bulkhead were painted with epoxy to seal the engine� XE "Engine" � compartment off from the rest of the fuselage. Finally, to seal the fuselage, two layers of ounce fiberglass at 45( were epoxied into the masking tape formed channels running the length of the fuselage on both sides.

With the fuselage� XE "Fuselage" � assembly completed, a number of tasks still remained. First, the tail cone had to be attached. Formica� XE "Formica" � templates for the front and back of the tail cone were constructed from Unigraphics drawings. Then they were bonded to the front and back faces of a 2” block of foam� XE "Foam" � which was then hot-wired down to the shape of the cone. Two layers of six ounce fiberglass� XE "Fiberglass" � were wrapped around the cone and allowed to cure. Then the foam was hollowed out, and Bondo� XE "Bondo" � was applied and sanded to smooth the surface. The tail cone was attached to the boom using epoxy� XE "Epoxy" �.

With the fuselage� XE "Fuselage" � structure complete, it was necessary to construct mounts� XE "Mounting" � to keep the hatches� XE "Hatches" � in place during flight. For the wing� XE "Wing" � mount hatch, four tabs with T-nuts on them were installed in the four corners of the hole for the hatch. Matching holes were also drilled on the hatch itself. This allowed the wing mount hatch to be held in place with four small screws. The other three hatches were mounted differently. The front two hatch mounts were simple pegs which fit into mounts in the fuselage. The pegs were actually nails which were installed into balsa� XE "Balsa" � and Formica� XE "Formica" � mounts. The mounts were epoxied� XE "Epoxy" � into the front of the hatch whose flange corners had been trimmed� XE "Trim" �. The mounts in the fuselage were also constructed of balsa and Formica and epoxied into the fuselage. The back two hatch mounts were spring boxes whose pegs fit into mounts in the fuselage. Each spring box was made from Formica for the sides, a nail secured inside a collar with a set screw for actuation, and a ball point pen spring. The back two hatch flange corners and slots for the set screw were Dremelled out and the spring boxes were all epoxied into the hatches. To secure all the hatches, epoxy and flocking� XE "Flocking" � fillets were made around them.

After extensive body work, which involved applying and sanding away Bondo� XE "Bondo" � for small holes in the skin� XE "Skin" � and epoxy� XE "Epoxy" � with microballoons� XE "Microballoons" � for larger imperfections, several holes had to be drilled in the fuselage� XE "Fuselage" �. These included holes in the boom for the ring-tail, on the fuselage bottom for the dolly� XE "Dolly" � mounts� XE "Mounting" �, and in the tail boom� XE "Tail boom" � for the ring-tail servo� XE "Servos" �. The final step in the fuselage construction was the skid� XE "Skids" � installation. For the tail skid, a slot was cut in the fuselage for the skid and the entire mechanism was simply epoxied and filleted into the fuselage. The main skids were more difficult. A drill guide was constructed from a length of wood with wall anchors installed at the correct locations and set at the correct angle for the drill. The drill guide was then clamped to the fuselage and the holes for the skids were drilled through the wall anchors. Finally, after bending the skids to the correct shape, they were installed in the holes and secured with Hysol bonding compound. And to further secure the back skids, a cross bar was bonded across them to maintain the correct angle between them.

Wing� XE "Wing" � Construction

The Raven utilized a balsa� XE "Balsa" � sheeted foam� XE "Foam" � core wing� XE "Wing" � with a carbon� XE "Carbon" � spar� XE "Spars" � on the quarter chord� XE "Chord" �. The leading and trailing edges along with the hinge coves were made of balsa, while the wing mounting� XE "Mounting" � parts and wing tip� XE "Wing tip" �s� XE "Wing tips" � were made of poplar� XE "Poplar" � plywood� XE "Plywood" �. A cross-sectional view� XE "section" � of the wing can be seen in figure 17.8.6.

Construction began with the fabrication of templates. Six templates were made for the wing� XE "Wing" � out of 1/64" plywood� XE "Plywood" �. These templates included two spar� XE "Spars" � templates and two templates for each the top and the bottom of the airfoil shaped foam� XE "Foam" � core of the wing (Figure 17.8.6). With the templates complete, the next step was to construct the spar. First two blocks of foam, each half the span of the wing, were hot wired� XE "Hot wire" � to the shape of the spar using the aforementioned templates. These blocks were then glued together to form a full length compression� XE "Compression" � mold� XE "Mold" � for spar lay up. The carbon� XE "Carbon" � was cut to allow the fibers to be oriented at 90(� XE "Wing" �. The carbon was coated with epoxy� XE "Epoxy" � and laid over the male part of the spar mold, then pressed into the desired shape by the female part. Weight� XE "Weight" � was place on top of the mold to maintain spar shape and to squeeze out any excess epoxy. After curing for 24 hours, the spar was cut from the mold and cleaned up by sanding.

While the spar� XE "Spars" � was curing, the foam� XE "Foam" � wing� XE "Wing" � cores were cut to shape. This involved using the templates to hot wire� XE "Hot wire" � the bottom first, then flipping the core over and hot wiring the top using another set of templates. This process leaves foam husks which are an outer shell of the airfoil shape. These are saved to aid in sheeting� XE "Sheeting" � and to protect the wing during construction. The wing cores were then cut in multiple places to allow the spar to fit in its location at the quarter chord� XE "Chord" � and to create conduits for airspeed bird and servos� XE "Servos" �.

When the foam� XE "Foam" � core was complete, the spar� XE "Spars" � was added. The two were joined by laying the spar in the bottom foam husks and epoxying it to the front and back parts of the wing� XE "Wing" �. Two 1/32" poplar� XE "Poplar" � ply ribs were also added near the wing root in this step to aid in transferring wing attachment loads� XE "Load" � to the spar. The top husk is put in place to hold everything in proper alignment, and the wing was allowed to cure for 24 hours.

Once the internal structure of the wing� XE "Wing" � was complete, the sheeting� XE "Sheeting" � process began. This involved joining 3/32" balsa� XE "Balsa" � sheets to make the required size for top and bottom wing sheeting. Epoxy� XE "Epoxy" � was applied to the back of the sheeting, and it was placed in the bottom husk. The core of the wing was placed on top, followed by the top sheeting and top husk. This was all weighted down to force the sheeting to bend to the proper shape and to prevent anything from moving during the cure process. After curing, the balsa leading and trailing edges were added. Theses edges were attached using wood glue, allowed to dry, and then shaped using X-acto knives and sand paper.

The semi-finished wing� XE "Wing" � was then ready for a layer of 0.75 ounce fiberglass� XE "Fiberglass" � to aid in painting. This was allowed to cure before applying another coat of epoxy� XE "Epoxy" � to fill in the weave of the glass cloth. After the final coat of epoxy cured, the wing was sanded until the tops of the fiberglass fibers were visible to reduce weight� XE "Weight" �.

After much sanding, the trailing edge surfaces were ready to be cut and finished. Lines were drawn on the wing� XE "Wing" � as dictated by the wing plan shown in figure 17.8.7, and the surfaces were cut out using X-acto knives. Balsa� XE "Balsa" � leading edges were added to the surfaces, and a corresponding balsa cove block was added to the trailing part of the wing. Both surfaces were sanded to fit each other and finished by capping the ends with 1/16" balsa and applying epoxy� XE "Epoxy" � resin� XE "Resin" � as a sealant. Hinges were then installed so that the rotation point was in the middle of the round in the surface leading edge.

Although the wing was near completion at this point, the flaperons� XE "Flaperons" � still needed to be mass balanced. This was done by gluing the flaperons to outside 3/4" of the cove. Next a one inch tab was cut from the wing� XE "Wing" � leaving a cut out in the wing for the mass balance� XE "Mass balance" �. The tab was now part of the flaperon and was capped with 1/16" balsa� XE "Balsa" � sides and a 1/4" rounded leading edge on the front. In order to determine the weight� XE "Weight" � required for mass balancing, the flaperon was supported about the hinge line in a manner that allowed it to rotate freely. Lead shot was then added to the front of the tab until the flaperon balanced. Finally the cut out in the wing was capped with balsa and sanded for a close fit with the tab. Again epoxy� XE "Epoxy" � resin� XE "Resin" � was used to seal any new balsa.

Now that the basic wing� XE "Wing" � structure was complete, equipment installation began. Servos� XE "Servos" � holes were cut of the bottom of the wing and were capped on the sides with balsa� XE "Balsa" �. Servo mounting� XE "Mounting" � rails were added as well. The wiring was then pulled through the conduits, and linkages� XE "Linkage" � were added� XE "IRON BIRD" �. Finally the servos were installed, and the surfaces were checked for proper movement. Servo fairings� XE "Fairing" �, shown in figure 17.8.8, were also constructed at this point to reduce drag� XE "Drag" � from the servos and to shield them from the elements. The servos were then removed until after painting.

The wing� XE "Wing" � was now ready to be fitted to the fuselage� XE "Mounting" �� XE "Fuselage" �. The front wing mount block assembly was cut out of 1/8” birch� XE "Birch" � ply and glued together. The wing sheeting� XE "Sheeting" � was cut to allow the mounting block to make contact with the main spar� XE "Spars" �. This assembly was then attached to the wing using flocking� XE "Flocking" � and epoxy� XE "Epoxy" �. The rear mounting birch ply plate was then glued to the wing. The wing was placed into position in the fuselage and holes were drilled for the rear mounting bolts and for the front carbon� XE "Carbon" � mounting dowels� XE "Dowel" �. A solid model of the wing installed in the fuselage is shown in figure 17.8.9. Following this last step, the wing was ready for final sanding, priming, and painting.

Tail

The ring-tail was constructed using the same basic foam� XE "Foam" � core and balsa� XE "Balsa" � sheeting� XE "Sheeting" � used in the construction of the wing� XE "Wing" �. This technique was used because of group familiarity with this type of construction. In order to make a cylindrical shape, however, 12 individual sections had to be constructed with edges cut to 15( and then later assembled to make a complete tail.

The construction began with foam� XE "Foam" � core templates for the tail. These templates, shown in figure 17.8.10, were screwed into three blocks of blue foam four inches wide by 18 inches long. These blocks were then hot-wired to the NACA 0010 shape of the tail to make 10 of the facets. Then the templates were attached to a block of high density foam which was sanded to the proper shape to create the two remaining facets. These facets were to be the ones located where the tail mounted to the aircraft. The four foam cores (three blue foam and one high density) were then sheeted with 1/16” balsa� XE "Balsa" � wood. Once this cured, the leading edge was attached using a 1/4" block of balsa and sanded to proper shape. The trailing edge was made by sandwiching a single uni-directional strip of carbon� XE "Carbon" � between two 1/8” pieces of balsa wood. The balsa-carbon-balsa 1/4" sandwich was then attached to the trailing edge of the foam core and sanded to the proper shape. A cross-section� XE "section" � of the tail airfoil is shown in Figure 17.8.11. When this step was finished there were three blue foam NACA 0010 wing� XE "Wing" � sections with a 18" span and a four inch chord� XE "Chord" � and one high density foam NACA 0010 wing section with a six inch span and a four inch chord.

The next step was construction of the fairings� XE "Fairing" � to cover the tail mounting� XE "Mounting" � rod. They were constructed from a 2.5” by six inch block of high density foam� XE "Foam" �. Templates� XE "Template" � of a NACA 0025 airfoil with a two inch chord� XE "Chord" � were attached to the foam and sanded to the correct shape. They were then covered with six ounce fiberglass� XE "Fiberglass" � for durability. A mistake was made while drilling the 3/8" hole for the mounting rod to slide through, so the fairings had to be reconstructed. The reconstruction process was exactly the same as the original construction technique. The single fairing was then cut into 2 three inch sections that could be sanded down to fit the tail boom� XE "Tail boom" �.

With the fairing� XE "Fairing" � construction complete the next step was the joining of the fairing to the mounting� XE "Mounting" � facets. First, the single high density foam� XE "Foam" � core NACA 0010 wing� XE "Wing" � was cut into two pieces which were beveled to the shape of a facet. Then an outline of the fairing shape was drawn on the facets. An X-acto knife was used to cut though the balsa� XE "Balsa" � wood and dig out approximately half of the high density foam core so the fairing would fit tightly into the facet. Before the fairings were attached to the facets, however, the mechanism for mounting the tail had to be incorporated. The mounting hardware in the tail consisted of a 1/2" x 1/2" x 1/4" block of plywood� XE "Plywood" � placed at one end of a fairing in the center of the 3/8" hole previously drilled for the mounting rod. This fairing and one mounting facet were then joined together. The other mounting facet was then flipped over and a one inch square as cut into the balsa wood sheeting� XE "Sheeting" �. This one inch square was then dug out until it met the airfoil cut out from the other side. A one inch square of 1/8” plywood was then glued into the facet to provide a hard surface for the hatch� XE "Hatches" �. The 3/8" hole for the rod was then drilled as well as a rectangular cut of 1/2" by 1/4" for the key to hold the rod in place. Finally the hatch was constructed from a 1 inch square of 1/8” ply with a 1/2" x 1/2" x 1/4" key glued to it. Then, with the hatch in place, the fairing was attached to the facet. Then the two mounting facets and the three blue foam core NACA 0010 wings were covered with fiberglass� XE "Fiberglass" � with 0.75 ounce fiberglass. After the glass dried an X-acto knife was used to cut the fiberglass around the hatch and the hatch was removed.

Now the mounting� XE "Mounting" � rods for the tail were constructed from a 3/8" diameter outer brass rod five inches long and a 11/32" diameter inner brass rod six inches . The inner rod had 1/2" long by 1/4" wide cut along the rod's axis on each end for the keys to fit in (Figure 17.8.12). This construction was used so the tail would be removable.

The next step in the construction of the tail was to cut the blue foam� XE "Foam" � core NACA 0010 wings� XE "Wing" � into 10 facets, with an outside length of 1.66 inches and an inside length of 1.45 inches. In order to make these facets fit together to form a ring, a band saw was tilted to a 15( angle which provided a symmetrical dodecagon when the facets were assembled. A six inch section� XE "section" � of blue foam was left over after finishing this step. A template� XE "Template" � of the tail was provided and the facets were placed in the template to ensure a symmetrical leading edge (Figure 17.8.13). The tail was then filled and sanded, primed, filled and sanded. With only painting left on the actual tail the inner rod was covered with tape and slid in place and the hatch� XE "Hatches" � attached with two small screws.

All mounting� XE "Mounting" � work for the tail was not quite finished. Inside the aircraft fuselage� XE "Fuselage" � two 1/2" blocks of wood were added to the tail boom� XE "Tail boom" � around the mounting holes for stiffness. The holes in the fuselage were then drilled, starting with pilot holes and gradually working up to a 3/8" bit. The wood was added for the set screw and that hole was also drilled. The tail was then mounted to the fuselage, at which point it was clear that the tail was on crooked. Corrective measures were taken and the tail appeared to be on the aircraft straight. The control horn� XE "Control horn" � was then added, and the tail was painted.

Painting

The final step before the plane was ready for flight testing� XE "Flight test" � was painting. Earlier in the semester, a red and white paint scheme for the aircraft was developed using Unigraphics� XE "Unigraphics" � and presented to Dr. Perkins for approval (Figure 17.8.14). After approval of the paint scheme and the completion of final body work touch-ups on the fuselage� XE "Fuselage" �, wing� XE "Wing" �, and ring-tail, the last coat of primer was sprayed on to each individual part to prepare them for painting. White paint was sprayed on to each part and allowed to dry for 24 hours (Figures 17.8.15 - 17.8.16). After the white paint was completely dry, areas of the aircraft that were to remain white were masked off according to the specifications of a dimensioned Unigraphics drawing of the approved paint scheme. When the masking was completed, a coat of red paint was applied. The red paint was allowed to dry for approximately four hours, and then the masking was pulled off to prevent the paint from chipping. After removing the tape, the aircraft was allowed to continue drying for approximately another 20 hours. When the paint was completely dry, it was ready for all other necessary work to get the aircraft ready for flight. The finished product is shown in figure 17.8.15.�Flight Testing

The flight test� XE "Flight test" � program of the Raven revealed a number of interesting points about the performance� XE "Performance" � of the design. The flight testing program consisted of four flight attempts. The following is a description of the conditions, points of interest, and results of each of these attempts. A summary of the takeoff� XE "Takeoff" � settings can be found in table 16.9.1.

Power Up Testing

Before attempting to launch the Raven, power up procedures were conducted in order to assure the proper operation of all systems. The aircraft was placed on the dolly� XE "Dolly" � in the exact configuration� XE "Configuration" � prescribed for takeoff� XE "Takeoff" �. This involved installing the required spacers, two extensions on all four posts and three washers on each wing� XE "Wing" � support, to give the desired takeoff incidence angle. The engine� XE "Engine" � was then started and run throughout its operating range� XE "Range" �. While the engine was operating, the control surfaces� XE "Control surfaces" � were deflected through their throw ranges. After testing all moving parts, the engine was run for a few more minutes. After engine shut down, the aircraft was inspected to be sure no parts had vibrated loose. This procedure also served as a practice run through the preflight inspection checklist.

First Flight Attempt

After ensuring that all systems were operational, the Raven was ready for the first flight attempt. This was conducted on Wednesday, 2 April 1997. Winds were relatively calm and skies were clear. After performing the preflight inspection at the NCSU Flight Research Facility in Butner, NC, the Raven was placed on the dolly� XE "Dolly" �. The dolly serves as a takeoff� XE "Takeoff" � platform for the Raven since this aircraft has no wheels. The dolly had been modified from its original configuration� XE "Configuration" � to allow two hard points on the fuselage� XE "Fuselage" � and two cradle points to support the wing as seen in figure 17.9.1� XE "Wing" �.

The dolly� XE "Dolly" � was attached to two bungee� XE "Bungee" � cords that were anchored on opposite sides of the runway. These bungee cords would provide the energy to propel the Raven and the dolly down the runway to achieve takeoff� XE "Takeoff" � speed. The dolly was then attached to a truck in order to pull it back to the appropriate takeoff distance (Figure 17.9.2), and a carabiner mechanism allowed the aircraft and dolly to be released at the appropriate time. Also, small pin was inserted in a hinge to connect the nose of the aircraft to the forward hard point. This pin was attached to the end of a cord, the length of which was equal to the distance necessary to reach takeoff speed. This cord was also attached to the truck. In this way, the pin was pulled out of the hinge at the appropriate time to allow the aircraft to lift off of the dolly.

Before launch, there were a few unanticipated issues to resolve regarding launch procedures, because a prior attempt to launch the Redeye aircraft had resulted in a mishap. Immediately upon release, the Redeye flipped off the dolly and landed upside down on the runway. This raised concerns about the dolly� XE "Dolly" � pull-back distance necessary for a successful launch. The pull-back distance was decreased from 50 feet to 42 feet as a result of the Redeye’s launch mishap. Also, because the release pin hinge was broken during the Redeye’s attempt, a third screw was installed on the Raven release pin hinge to ensure the integrity of the hinge under the high loads� XE "Load" � of takeoff� XE "Takeoff" �.

Upon lift-off from the dolly� XE "Dolly" �, the Raven assumed a strong nose up attitude as seen in figure 17.9.3. This was corrected by down elevator. In addition� XE "Takeoff" �, there was significant yawing� XE "Yaw" � to the right. This continued until the Raven yawed nearly 180(, almost facing the opposite direction from takeoff. The aircraft then rolled to an inverted position and the pilot lost control. The engine� XE "Engine" � was shut down in anticipation of the crash. The right wing� XE "Wing" � was the first point of impact, followed by the nose. Total time from release to impact was approximately 11.14 seconds.

The damage sustained in this attempt was significant. The wing, shown in figure 17.9.4,� XE "Wing" � was broken at the root. Four of the 12 facets of the ring-tail were damaged (Figure 17.9.5), and the engine� XE "Engine" � mount� XE "Mounting" � was broken. Also, the fuselage� XE "Fuselage" � skin� XE "Skin" � was pierced and split at a number of points near the nose (Figure 17.9.6). The fuselage and tail were salvageable, but the wing required a complete rebuild.

There were a number of theories presented to account for the catastrophic ending to the first flight attempt. Attention was first given to a significant approximation in the aerodynamic� XE "Aerodynamics" � model. The thickness of the wing� XE "Wing" � had been chosen such that the flaperon� XE "Flaperons" � and deflector� XE "Deflectors" � servos� XE "Servos" � protruded from the lower surface of the wing (Figure 17.8.8). Fairings� XE "Fairing" � were constructed to cover these protrusions in order to minimize their effects on the flow under the wing. These servos and fairings were left off of the aerodynamic model of the wing, assuming they would not affect the performance� XE "Performance" � of the aircraft significantly. This was not necessarily the case, however, and these fairings were one possible explanation for the crash.

The second theory also involved a difference between the aerodynamic� XE "Aerodynamics" � model and the actual flight scenario. PMARC� XE "PMARC" � could not simulate the flow of the propwash over the aircraft, so it is conceivable that the resulting power effects changed the stability� XE "Stability" � characteristics of the Raven enough to result in the behavior observed in this flight attempt.

Finally, it was proposed that the tail employed on the Raven was not large enough to provide the stabilizing forces necessary for successful flight. If this theory were correct, the chord� XE "Chord" � and diameter of the ring could be increased to provide a more stabilizing tail, with the diameter being the most significant variable. It was also possible that if the tail were not mounted perfectly straight, this would initiate a yaw� XE "Yaw" � on takeoff� XE "Takeoff" � regardless of the size of the tail. Therefore, no matter what tail sized was used, it was clear that care should be taken in the future to ensure the tail was mounted correctly.

The only change that was made during the repair process was to change the wing� XE "Wing" � design so that the servos� XE "Servos" � fit almost completely within the wing. The single spar� XE "Spars" � design was altered to a two-spar system allowing the servos to move forward into a thicker part of the airfoil. These changes are discussed in detail in Section 10.

Second Flight Attempt

A second flight attempt was conducted on Tuesday, 22 April 1997. On this attempt, the launch procedure was similar to the previous attempt except that the pull-back distance was arbitrarily increased by approximately five feet. Upon launch, the aircraft did not yaw� XE "Yaw" � significantly in either direction. However, it did demonstrate a radical pitching� XE "Pitch" � up maneuver, at which point the aircraft departed and rotated 180( around the y-axis. The aircraft impacted the ground directly on the nose approximately eight seconds after dolly� XE "Dolly" � release.

The damage was more severe on the fuselage� XE "Fuselage" � than the first attempt but still repairable, and fortunately the tail and wing� XE "Wing" � were spared. With no clear explanation for the crash, it was decided that mounting� XE "Mounting" � a larger tail on the Raven might solve the problem of pitch� XE "Pitch" � control on takeoff� XE "Takeoff" �. Therefore, a new tail was constructed with a chord� XE "Chord" � of six inches and a diameter of 6.75 inches. Again these changes are covered in detail in Section 10.

Third Flight Attempt

A third flight was attempted on Thursday, 24 April 1997. This was the first successful flight of the Raven, lasting approximately 8 minutes and 53 seconds. It was noted that the aircraft was sensitive to pilot commands, as slight movements of the transmitter controls were necessary to make drastic changes in the attitude of the aircraft around all axes. As predicted from dynamic simulation results, the aircraft’s Phugoid mode was easily exited in pitch� XE "Pitch" �. The pilot was constantly adjusting the pitch to account for this during most o