INTRODUCTION
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Final Report (Volume II)
Raven Design Team
Brian Hucks Group Leader
Rachel Block Structures
Ryan Czerwiec Aerodynamics
Elton Fairless Internal Systems & Procurement
Dustin Keck Report Organizer
Ryan McDaniel Performance
Matt Shaver Stability & Controls
Michael Sirmans Weights & Balances
Sean Zamora CAD & Unigraphics
AE Senior Design Project (Group II)
Advisor: Dr. John N. Perkins
Department of Mechanical & Aerospace Engineering
NCSU
Raleigh, NC
5 May 1997
ABSTRACT
The need for short-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 as a predecessor to this tube-launched solution. Issues of control surface configuration and sizing were central to the success of this project. Specifically, the ring-tail sizing was integral to stability. Also, issues of yaw control were addressed extensively. Molded composite construction methods were utilized on the fuselage and spars while foam core and balsa sheeting techniques were used on the wing and tail. The Raven aircraft was constructed on time and under budget, with the first flight attempt on 2 April 1997. The first successful flight occurred on 24 April 1997.
Table of Contents
1. Introduction 7
2. Design History 8
2.1 Wing 8
2.2 Tail 9
2.3 Skids 9
3. Wind Tunnel Testing 10
4. Stability and Controls 11
4.1 Drag Rudder to Deflectors 11
4.2 Controller 11
4.3 Trim 12
5. Structures 13
5.1 Structures Review 13
5.2 Finite Element Analysis 13
5.3 Conclusions 15
6. Weights and Balances 16
7. Performance 17
8. Construction 18
8.1 Fuselage 18
8.1.1 Plug 18
8.2 Wing Construction 22
8.3 Tail 24
8.4 Painting 26
9. Flight Testing 27
9.1 Power Up Testing 27
9.2 First Flight Attempt 27
9.3 Second Flight Attempt 29
9.4 Third Flight Attempt 29
9.5 Fourth Flight Attempt 30
10. Modifications and Repairs 31
10.1 First Flight Attempt 31
10.2 Second Flight Attempt 32
10.3 Future Modifications 33
11. Budget Overview 34
12. Concluding remarks 35
12.1 Conclusions 35
12.2 Future Research 35
13. Acknowledgments 36
14. References 37
15. Appendix 38
15.1 Program Trim 38
16. Tables 41
16.1 Introduction 41
16.2 Design History 41
16.3 Wind Tunnel Testing 42
16.3.1 New Pitch Rate and Yaw Rate Coefficients 43
16.4 Stability and Controls 44
16.4.1 Stability Derivatives for 4” Tail Chord 45
16.4.2 Trim Angles for Varying Flight Configurations 46
16.5 Structures 47
16.5.1 Material Properties 48
16.5.2 Summary of ANSYS Results 49
16.6 Weights and Balances 50
16.6.1 Predicted vs. Actual Weights 51
16.6.2 Weight and Balance Summary 52
16.7 Performance 53
16.7.1 Parameter Values 54
16.7.2 Performance Summary 55
16.8 Construction 55
16.9 Flight Testing 56
16.9.1 Takeoff Settings for Individual Flights 57
16.10 Modifications and Repair 58
16.10.1 Effects of Vertical Fins 59
16.10.2 Effect of Servo Boxes 60
16.10.3 Aerodynamic Coefficients of Final Configuration 61
16.10.4 Stability Derivatives 6” Tail Chord 62
16.11 Budget Overview 63
16.11.1 Initial and Repair Budgets 64
17. Figures 65
17.1 Int 65
17.2 Design History 66
17.2.1 Overall Configuration with 4” Tail 67
17.2.2 Overall Configuration with 6” Tail 68
17.3 Wind Tunnel Testing 69
17.3.1 Wind Tunnel Testing Setup 70
17.3.2 Aircraft Mesh with Deflectors and Resized Flaperons 71
17.3.3 Pressure Contour with Deflector 72
17.3.4 Pressure Contour with Ailerons 73
17.4 Stability and Controls 74
17.4.1 Aileron and Deflector Comparison for 0.1 rad Doublet 75
17.4.2 Block Diagram of Yaw Damper 76
17.4.3 Yaw Rate Comparison for Baseline and Yaw Controller 77
17.4.4 Block Diagram of Non-linear Servo 78
17.5 Structures 79
17.5.1 Fuselage Internal Structure 80
17.5.2 Finite Element Mesh of Wing 81
17.5.3 Wing Tip Deflection at ( = 0(, 1g Loading 82
17.5.4 Wing Tip Deflection at ( = 0(, 6g Loading 83
17.5.5 Wing Tip Deflection at ( = 5(, 1g Loading 84
17.5.6 Wing Tip Deflection at ( = 5(, 6g Loading 85
17.5.7 Wing Tip Deflection with (f = 5( 86
17.5.8 Wing Tip Deflection with (f = 30( 87
17.5.9 (y at ( = 0(, 1g Loading 88
17.5.10 (y at ( = 0(, 6g Loading 89
17.5.11 (y at ( = 5(, 1g Loading 90
17.5.12 (y at ( = 5(, 6g Loading 91
17.5.13 (y at (f = 5( 92
17.5.14 (y at (f = 30( 93
17.5.15 Von Mises Stress at ( = 0(, 1g Loading 94
17.5.16 Von Mises Stress at ( = 0(, 6g Loading 95
17.5.17 Von Mises Stress at ( = 5(, 1g Loading 96
17.5.18 Von Mises Stress at ( = 0(, 6g Loading 97
17.5.19 Von Mises Stress at (f = 5( 98
17.5.20 Von Mises Stress at at (f = 30( 99
17.6 Weights and Balances 100
17.6.1 Internal Component Layout 101
17.7 102
17.8 Construction 102
17.8.1 Sample Fuselage Cross-Section 103
17.8.2 Completed Templates and Cross-Section Layout 104
17.8.3 Templates Mounted on Blue Foam 104
17.8.4 Joined Plug 105
17.8.5 Top Half of Mold 106
17.8.6 Original Single Spar Wing Cross-Section 107
17.8.7 Wing Layout 108
17.8.8 Wing Servo Fairings 109
17.8.9 Wing Mounting System 110
17.8.10 Ring-Tail Foam Core Template 111
17.8.11 Ring-Tail Facet Cross-Section 112
17.8.12 Ring-Tail Mounting Mechanism 113
17.8.13 Ring-Tail Facets in Assembly Template 114
17.8.14 Paint Scheme 115
17.8.15 Fuselage After Coat of White Paint 116
17.8.16 Wing After Coat of White Paint 117
17.8.17 Completed Raven Aircraft 118
17.9 Flight Testing 119
17.9.1 Launch Dolly 120
17.9.2 Aircraft Prepared for Takeoff 121
17.9.3 First Flight Attempt Immediately After Takeoff 122
17.9.4 Wing Damage Sustained After First Flight Attempt 123
17.9.5 Ring-Tail Damage Sustained After First Flight Attempt 124
17.9.6 Fuselage Damage Sustained After First Flight Attempt 125
17.10 Modifications and Repair 126
17.10.1 Exploded View of Wing 127
17.10.2 Pressure Contour of Sideslip with Fins 128
17.10.3 Aircraft Mesh with Servo Fairings 129
17.10.4 Baseline Pressure Contour with Servo Fairings 130
17.10.5 Baseline Pressure Contour with 6” Tail 131
17.10.6 Ground Effect Survey 132
17.10.7 Pitch Angle Comparison for 0.1 rad Doublet 133
17.10.8 Pitch Rate Comparison for 0.1 rad Doublet 134
17.10.9 Yaw Angle Comparison of Ailerons for 0.1 rad Doublet 135
17.10.10 Yaw Angle Comparison of Deflectors for 0.1 rad Doublet 136
17.10.11 Roll Angle Comparison of Ailerons for 0.1 rad Doublet 137
17.10.12 Roll Angle Comparison of Deflectors for 0.1 rad Doublet 138
17.10.13 Wing Pivot Mechanism 139
18. Index 140
1 Introduction
The 1996-97 Senior Design Project in aerospace engineering at North Carolina State University was the design, construction, and flight testing 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
• Maximum stowed radial dimension of 7”
• Maximum 60” wing span
• Minimal weight
• Design loads of +4 and -2
• Safety factor of 1.5
• 10% power-off static margin
• OS Max 61SFABC-P engine/tractor propeller
• Swing-wing upgrade capability (for tube-launching)
• Dolly launch with skid 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 was developed early on, with the flexibility to allow changes when necessary. During flight testing, this flexibility in design would prove very important. This configuration was based on a commitment to meet or exceed all of the design criteria.
2 Design History
At the end of the fall semester, the overall 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 was an SD7062 airfoil with a 6.4 inch chord and five foot wing span. The control surfaces on the wing consisted of flaperons for roll control and drag rudders for 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. 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.
1 Wing
At the end of the fall semester, PMARC analysis revealed that the 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 force to our advantage through the use of control surfaces called deflectors. Stability and controls analysis in the fall had already determined that the drag rudders would not provide sufficient yaw control for the aircraft. The deflectors were small, inboard flaps that deflected in the up direction only. For example, an upwards 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 control without interfering with yaw control.
The wing also underwent a structural evolution in the spring semester. Originally, the wing spar was to be made out of wood. However, little analysis was done on this 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 to be too high for a wood spar at the upper limit load of +6g. Therefore, a strong, yet lightweight alternative was needed. This alternative was a C-channel carbon fiber spar placed at the quarter 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 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 were also constructed to help reduce the 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.
2 Tail
Although the tail was the one thing that remained a constant throughout the construction process, this was not the case during flight testing. After the wing servo fairings 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.
3 Skids
During the third and fourth flights, the original main skid design proved to be unable to sustain the landing loads, which were estimated to be much higher than the design load 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 tip skids and the tail skids withstood all flights and were therefore left unchanged.
3 Wind Tunnel Testing
The first aerodynamic analysis in the spring semester involved an investigation of the adverse aileron yaw, which is described in detail in volume 1. Wind tunnel testing was conducted using a full-size fuselage and a 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, the wind tunnel test showed that the adverse yaw originated from vortices on the inboard edges of the ailerons. Thus, by moving the ailerons outboard and using small inboard flaps, called deflectors, 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 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 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 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 described in the design modifications 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 analyses for pitch and 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.
4 Stability and Controls
1 Drag Rudder to Deflectors
As discussed in Section 3, the ineffective drag rudders were replaced with deflectors. The deflectors were modeled in 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. 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 when the deflector was used to yaw the aircraft. It had little response to the deflector in yaw, but a roll response was clearly seen.
2 Controller
Another requirement for the project was to design a 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 for this particular aircraft. While each person worked on identical controllers, the amount of damping on the Dutch 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, a continuous time system was developed which included aircraft dynamics, servo, 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. 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 was designed using a continuous time non-linear servo 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, Cn(, Cnr, and Clp were the most sensitive to changes. Noise was also added into the 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.
3 Trim
The trim angles for takeoff and level flight were calculated using a self-written program found in the Appendix. The equations used were for linear lift and moment curves and included all control surface deflections and incidences. The settings for different flight configurations are shown in Table 16.4.2.
5 Structures
1 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.
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 skin was to be made of five layers of bi-directional Kevlar. This was reduced to only three layers to decrease weight. Additionally, the original stringers in the aircraft were to be made of strips of poplar wood. Instead, the stringers were made of uni-directional 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. Originally, the wing had been designed with a single wooden spar located at the quarter chord. But preliminary analysis in the spring semester revealed that this spar would be unable to withstand applied loads 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 fiber laid at 90( located again at the quarter chord. This was the configuration of the wing for the first flight. However, as a result of a modification made during flight testing, the spar had to be moved forward to the leading edge of the wing (see 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.
2 Finite Element Analysis
The two primary tools used in the analysis were Unigraphics and ANSYS, both described in Volume 1. Because the wing is symmetric about the aircraft centerline, it was only necessary to model half the span. The wing mesh, shown in figure 17.5.2, is basically a balsa wood shell with two carbon fiber spars running through it at the leading and trailing edges (see table 16.5.1 for material properties). The 2480 elements in this mesh are spaced closer together at the wing root because in these areas the stresses 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 at the top, and the leading edge is on the left.
The wing was analyzed at 50 mph under both 1g and 6g loads for ( = 0( and 5(, as well as with 5( and 30( flap deflections. 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 system.
In the analysis, the focus was on both the wing tip deflection, (y (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 was not closely examined due to the fact that ANSYS has some problems with Poisson’s Ratio for 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 tip deflections 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 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 causes compressive and tensile stresses 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 while the lower surface is in 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 at the wing root. This is the area on the spar that is connected in with the wing mount 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 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 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 fiber are both 180,000 psi. One might also notice that like the wing tip deflection results, the numbers for the 30( flap 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 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 can handle the applied loads. 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 was added to the 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 deflection case.
3 Conclusions
Based on the analysis shown here, the two-spar wing is sufficient to withstand the applied aerodynamic loads with the current wing mount system. The highest 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. This analysis is supported through some qualitative flight test information. During flight testing, there have been several landings 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 in the wing as well as the fiberglass on the outside of the wing, which makes the analysis conservative from the start.
6 Weights and Balances
At the end of last semester the estimated dry aircraft 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, paint, and wiring. Another reason was the loss of almost 1.2 lbs from the lack of two digital cards and the flight data recorder (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 and FDR were not in the aircraft at first flight required other components to be moved to achieve the required 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 (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 before the first flight to balance the plane longitudinally. The component layout used to reach the required static margin, minus the ballast weight, is shown in figure 17.6.1. No weight was required for lateral balance.
The aircraft dry weight was 9.5 lbs after reconstruction of the wing. The weight reduction was attributed to the loss of fiberglass and 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 due to repair work. Once again, no ballast weight was needed laterally.
With the introduction of the new tail and more resin and fiberglass to the front of the fuselage the aircraft dry 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 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 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.
7 Performance
Analysis this semester involved recalculating the aircraft performance characteristics (takeoff, cruise, and landing) with the new 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 calculations were performed. The actual control surface deflections for various flight conditions are presented in Section 4.
Takeoff calculations began by re-running the Bungee program to determine the bungee pullback distance with the updated weight. To reach the takeoff 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 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. In order to achieve this condition, the wing is mounted on the 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. 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 will no longer be zero.
Finally, the original landing calculations were performed by a FORTRAN program which calculated gliding performance at various velocities using the non-dimensional CL and CD alone. Therefore, the aerodynamic coefficients at the actual aircraft 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, cruise, and landing performance are summarized in table 16.7.2.
8 Construction
1 Fuselage
1 Plug
The first step in the fuselage construction was the formation of the 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. Hard copies of the cross-sections were then cut in half along the horizontal center line. Only half of the template was required, because except for the wing mount 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 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 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 and ultimately to the finished fuselage skin. The finished templates and the cross-section layout are shown in Figure 17.8.2.
The next step in the plug construction was to fabricate the blue 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 1Ó and 2Ó apart, but to account for the 1/32Ó thickness of the Formica templates, the foam blocks had to be hot wired 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 blocks finished, assembly of the 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 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 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 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 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 was found later.
After completing the plug, mold fabrication began. First, a cradle to hold the plug was constructed. The shape of bottom half of the plugÕs constant area section was hot wired out of the top of a length of blue 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. 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. 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 would release easily. Then the actual mold was ready to be laid up. First, two layers of nine ounce fiberglass were laid up over the bottom half of the 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 skin in. The finished top half of the mold is presented in Figure 17.8.5.
Before laying up the fuselage skin, reference lines had to be drawn in the mold to mark the location of the stringers and hatches. References lines parallel to the fuselageÕs constant area section were drawn on the ledges of the mold. A tool was constructed of foam and 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, it was decided to join the two halves of the fuselage skin on the outside with strips of 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 prepared, the lay up of the bottom half of the fuselage began. Three layers of 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 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 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 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 fiber 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 are on the upper surface of the aircraft, they had to be constructed first so that the fuselage skin could be laid up on top of them. The front and back positions of the hatches were marked on the mold ledges. The sides of all hatches were at 45( from vertical. Once again, the foam-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 strips were pressed into the inside corners of the hatches which were then filled with a mixture of epoxy and microballoons. When completed, this would form the sharp edges of the hatches. This was necessary because the 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 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. 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 was laid up identically to the bottom half except holes had to be pre-cut to allow the Kevlar to fit over the 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 construction, bulkhead fabrication, actually began while the skin was being laid up. Like the fuselage templates, Unigraphics printouts were bonded to poplar plywood which was then roughly cut to shape. And like the hatches, the bulkheads were located in the fuselage by measuring from the nose and marking the ledge of the 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 and wing mount bulkheads were glassed with a single layer of six ounce fiberglass and vacuum bagged. Also during this time, all other internal structures such as the SAS mount and skid hardpoints were installed. The final installation involved bonding the bulkheads first into the bottom half of the fuselage using epoxy and 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 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 assembly completed, a number of tasks still remained. First, the tail cone had to be attached. 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 which was then hot-wired down to the shape of the cone. Two layers of six ounce fiberglass were wrapped around the cone and allowed to cure. Then the foam was hollowed out, and Bondo was applied and sanded to smooth the surface. The tail cone was attached to the boom using epoxy.
With the fuselage structure complete, it was necessary to construct mounts to keep the hatches in place during flight. For the 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 and Formica mounts. The mounts were epoxied into the front of the hatch whose flange corners had been trimmed. 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 fillets were made around them.
After extensive body work, which involved applying and sanding away Bondo for small holes in the skin and epoxy with microballoons for larger imperfections, several holes had to be drilled in the fuselage. These included holes in the boom for the ring-tail, on the fuselage bottom for the dolly mounts, and in the tail boom for the ring-tail servo. The final step in the fuselage construction was the skid 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.
2 Wing Construction
The Raven utilized a balsa sheeted foam core wing with a carbon spar on the quarter chord. The leading and trailing edges along with the hinge coves were made of balsa, while the wing mounting parts and wing tips were made of poplar plywood. A cross-sectional view 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 out of 1/64" plywood. These templates included two spar templates and two templates for each the top and the bottom of the airfoil shaped 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 to the shape of the spar using the aforementioned templates. These blocks were then glued together to form a full length compression mold for spar lay up. The carbon was cut to allow the fibers to be oriented at 90(. The carbon was coated with epoxy and laid over the male part of the spar mold, then pressed into the desired shape by the female part. 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 was curing, the foam wing cores were cut to shape. This involved using the templates to 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 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 and to create conduits for airspeed bird and servos.
When the foam core was complete, the spar 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. Two 1/32" poplar ply ribs were also added near the wing root in this step to aid in transferring wing attachment loads 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 was complete, the sheeting process began. This involved joining 3/32" balsa sheets to make the required size for top and bottom wing sheeting. 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 was then ready for a layer of 0.75 ounce fiberglass to aid in painting. This was allowed to cure before applying another coat of 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.
After much sanding, the trailing edge surfaces were ready to be cut and finished. Lines were drawn on the wing as dictated by the wing plan shown in figure 17.8.7, and the surfaces were cut out using X-acto knives. 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 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 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 leaving a cut out in the wing for the mass balance. The tab was now part of the flaperon and was capped with 1/16" balsa sides and a 1/4" rounded leading edge on the front. In order to determine the 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 resin was used to seal any new balsa.
Now that the basic wing structure was complete, equipment installation began. Servos holes were cut of the bottom of the wing and were capped on the sides with balsa. Servo mounting rails were added as well. The wiring was then pulled through the conduits, and linkages were added. Finally the servos were installed, and the surfaces were checked for proper movement. Servo fairings, shown in figure 17.8.8, were also constructed at this point to reduce drag from the servos and to shield them from the elements. The servos were then removed until after painting.
The wing was now ready to be fitted to the fuselage. The front wing mount block assembly was cut out of 1/8” birch ply and glued together. The wing sheeting was cut to allow the mounting block to make contact with the main spar. This assembly was then attached to the wing using flocking and 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 mounting dowels. 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.
3 Tail
The ring-tail was constructed using the same basic foam core and balsa sheeting used in the construction of the 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 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 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 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 of the tail airfoil is shown in Figure 17.8.11. When this step was finished there were three blue foam NACA 0010 wing sections with a 18" span and a four inch 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 to cover the tail mounting rod. They were constructed from a 2.5” by six inch block of high density foam. Templates of a NACA 0025 airfoil with a two inch chord were attached to the foam and sanded to the correct shape. They were then covered with six ounce 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.
With the fairing construction complete the next step was the joining of the fairing to the mounting facets. First, the single high density foam core NACA 0010 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 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 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. 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. 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 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 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 core NACA 0010 wings 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 of blue foam was left over after finishing this step. A 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 attached with two small screws.
All mounting work for the tail was not quite finished. Inside the aircraft fuselage two 1/2" blocks of wood were added to the 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 was then added, and the tail was painted.
4 Painting
The final step before the plane was ready for flight testing was painting. Earlier in the semester, a red and white paint scheme for the aircraft was developed using 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, 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.
9 Flight Testing
The flight test program of the Raven revealed a number of interesting points about the 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 settings can be found in table 16.9.1.
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 in the exact configuration prescribed for takeoff. This involved installing the required spacers, two extensions on all four posts and three washers on each wing support, to give the desired takeoff incidence angle. The engine was then started and run throughout its operating range. While the engine was operating, the 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.
2 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. The dolly serves as a takeoff platform for the Raven since this aircraft has no wheels. The dolly had been modified from its original configuration to allow two hard points on the fuselage and two cradle points to support the wing as seen in figure 17.9.1.
The dolly was attached to two 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 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 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 of takeoff.
Upon lift-off from the dolly, the Raven assumed a strong nose up attitude as seen in figure 17.9.3. This was corrected by down elevator. In addition, there was significant yawing 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 was shut down in anticipation of the crash. The right 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, was broken at the root. Four of the 12 facets of the ring-tail were damaged (Figure 17.9.5), and the engine mount was broken. Also, the fuselage 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 model. The thickness of the wing had been chosen such that the flaperon and deflector servos protruded from the lower surface of the wing (Figure 17.8.8). Fairings 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 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 model and the actual flight scenario. PMARC could not simulate the flow of the propwash over the aircraft, so it is conceivable that the resulting power effects changed the 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 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 on 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 design so that the servos fit almost completely within the wing. The single spar 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.
3 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 significantly in either direction. However, it did demonstrate a radical pitching 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 release.
The damage was more severe on the fuselage than the first attempt but still repairable, and fortunately the tail and wing were spared. With no clear explanation for the crash, it was decided that mounting a larger tail on the Raven might solve the problem of pitch control on takeoff. Therefore, a new tail was constructed with a chord of six inches and a diameter of 6.75 inches. Again these changes are covered in detail in Section 10.
4 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. The pilot was constantly adjusting the pitch to account for this during most of the flight. The deflectors were also tested to determine their effectiveness in yaw control. Once deflected, they caused the aircraft to roll and pitch down. This flight is evidence that the deflectors are more powerful in roll than they are in yaw.
The first landing attempt during the flight was aborted due to a lack of significant rudder control. Turning from base onto final approach was difficult, and the aircraft had too much altitude and speed to make a safe landing. The second landing attempt was a successful, but very hard landing, possibly over the 6g load limit. Fortunately, the only damage sustained was that the nose skid broke free from the aircraft and was lost in the grass. The skid was replaced overnight, and the aircraft was ready for flight the following day.
During this flight, however, the new tail proved only marginally stable. The aircraft still appears to have many strange power-on effects. These effects, the reasons for the failure of the original tail, and the failure of the deflectors to produce yaw will be investigated in the future.
5 Fourth Flight Attempt
On Friday, 25 April 1997, the aircraft had it’s second successful flight. The aircraft seemed to have improved in the Phugoid mode because it was not pitching as much as in the previous flight. This could be attributed to the calmer winds, as there had been some wind gusts on the previous day. This attempt resulted in the longest flight of either group to date, lasting over 10 minutes. As in the previous flight, three landing attempts were made before the aircraft finally touched down on the fourth try. The front skid once again collapsed on landing. The flight was otherwise a complete success.
10 Modifications and Repairs
1 First Flight Attempt
During the first flight attempt, the original Raven wing was destroyed (Figure 17.9.4). One of the theories as to the cause of this crash centered around possible flow disturbances created by the wing servo fairings (Figure 17.8.8). Since construction of a new wing was necessary, changes were made to allow the servos to be buried in the wing and eliminate the need for the fairings.
In order to accommodate the wing servos, the spar arrangement had to be altered. There are now two spars in the wing: one located right behind the balsa leading edge and the other at the rear of the wing. The trailing edge spar actually forms the cove for the trailing edge surfaces. The construction method was similar to the first wing, but the spars were laid up after the wing had been sheeted by digging out foam from the front and back of the balsa sheeted wing where the spar was located. The spar was then laid up into these coves and allowed to cure.
In order to speed finishing of the wing after the spars were complete, the wing surfaces was monokoted wing as opposed being covered with fiberglass and painted. Monokot is a colored plastic film that is applied using an iron. This reduced construction time by removing the priming and some of the sanding necessary for the wing construction, as well as drying time associated with painting. This new configuration is shown in as an exploded view in figure 17.10.1.
During the first attempt at flight the tail sustained damage to four facets (Figure 17.9.5). Replacement facets were made from the remainder of the blue foam core. When the tail was placed on the aircraft again it was once again noted that the tail wasn't on the aircraft straight. An enormous amount of effort were made to ensure the tail was straight for the second flight.
As previously discussed, the aircraft appeared to be unstable in yaw during the first flight attempt. Several explanations arose for this behavior (see Section 9). In an effort to counteract the aircraft's seemingly poor yaw stability, no matter what the cause, vertical fins were added to the aerodynamic model. Figure 17.10.2 exhibits the pressure contour of sideslip flow with the addition of fins in order to observe the change in directional stability. As seen in table 16.10.1 , the fins actually reduce lateral stability, because they substantially reduce the contribution of the ring while not performing efficiently themselves. They likely spill flow around the ring and partially turn the flow through the ring so that it experiences little sideslip. As such, the idea of adding vertical fins as stabilizers was abandoned.
In order to account for the effects of the servo fairings, the wind tunnel model was fitted with fairings and tested. Wind tunnel analysis showed that the fairings had no effect on the side force or yawing moment of the tail. This conclusion was verified by adding fairings to the aerodynamic model, as seen in the surface mesh of figure 17.10.3. To allow for downward movement of the flaperons, the outboard fairings required a blunt rear face. Because PMARC cannot simulate viscous separation, its solution is not entirely accurate for these blunt faces, as shown by somewhat unusual flow patterns over them, given in figure 17.10.4. However, the effect is small and the results at least show the aerodynamic trends caused by the fairings. Table 16.10.2 gives a comparison that shows the changes caused by the fairings. PMARC agrees with the wind tunnel data in that it shows little change in the directional stability of the aircraft. In fact, what little change exists actually seems to stabilize the configuration more. The most notable effects are the reduction in lift and a doubling in the tendency to roll with sideslip. The lateral derivatives increased in value.
2 Second Flight Attempt
During the second flight attempt, the aircraft appeared to have problems in pitch rather than yaw. The group decided to try a larger tail, largely because the Redeye aircraft, having a larger tail, had flown successfully. The diameter and chord of the tail were increased to the same size as the Redeye tail. The new tail was symmetrical with a six inch chord and a 6.75 inch outer diameter. The tail was shaped on a lathe, sanded, and then covered with six ounce fiberglass. New fairings were constructed as well as a new mounting rod. The mounting system for the new tail was constructed in almost the same fashion as before with the hinge point again at the quarter chord of the tail. The only difference between the mounting was in the fairing attachment to the tail, with the new fairings only glued to the tail surface and not embedded in the facet.
Before any flights with the new tail were attempted, it was added to the aerodynamic model, and all of the analysis was repeated. Figure 17.10.5 shows the baseline pressure contour as an example of the new configuration. The new tail exhibits the same suction on each side of the inside of the ring as the old tail, but this suction is larger in magnitude, which may explain the added stability of the new tail. A larger symmetric, stabilizing force is more sluggish to respond to disturbances. Table 16.10.3 displays the aerodynamic coefficients for the final configuration. In addition to the standard aerodynamic cases, a ground effect survey was performed, the results of which are provided in figure 17.10.6.
From PMARC data, a simulation was conducted to compare the dynamic responses of the aircraft with both tails. This comparison is shown in figures 17.10.7 through 17.10.12. Figures 17.10.7 and 17.10.8 represent the longitudinal responses of the tails. Both pitch and pitch rate responses were decreased by the larger tail. Although the response decreased, there was still enough control authority in flight. The main reason the tail size was increased was to increase the lateral stiffness of the aircraft. Figures 17.10.9 and 17.10.10 show that yaw responses decreased for deflectors and ailerons. The roll response in figures 17.10.11 and 17.10.12 shows that changing the size of the tail does not change the roll characteristics of the aircraft. From the data, some important items to note were that the weathercock stability increased and many longitudinal derivatives increased due to more flow going through the larger tail. This can be seen in tables 16.4.1 and 16.10.4. This change in the aircraft design did prove air worthy because the third flight was a success.
3 Future Modifications
At the outset of this project, the conceptual design was set forth as a tube-launched reconnaissance vehicle that was easily transportable and required the smallest possible diameter tube. This was kept in mind in the design of the wing, which was positioned on the aircraft so that it could pivot to line up with the length of the fuselage. Obviously the wing on this aircraft does not pivot; however, a mechanism has been designed that would provide that capability. Figure 17.10.13 shows this mechanism in detail. The two longitudinal support cross-bars carry the pivoting mechanism. A lateral support cross-bar serves as the mounting plate. A spring is attached to this plate and to a plate inside the wing. The plate in the wing distributes the loads to the fore and aft wing spars. In order to launch the aircraft, the wing would be lifted out of its cradle in the fuselage and pivoted parallel to the fuselage. This action would stress the spring axially and in torsion. Therefore, the spring would provide the force necessary to swing the wing to flight position and pull it down into the fuselage cradle upon launch release. This release is accomplished with a servo. When the wing is pulled into the cradle, spring-loaded carbon aeroshafts slide into place to hold the wing at two points on the leading and trailing edges. This locks the wing into the position it was in during the Raven flight testing described in this project.
In addition to pivoting the wing, the engine would be changed from a gasoline-powered engine to an electric engine and fitted with a collapsible propeller. All of the skids would also have to be deployable to allow them to fit in the tube for storage and launch.
11 Budget Overview
The budget for this project was $750.00. This was to include all construction items, parts, and reports for the Raven. Items not included in this budget were all composite materials, epoxy/resin, paint, electronic components (radio receiver/transmitter, FDR, SAS, digital cards, battery packs), servos, and the engine. In the fall semester $265.54 was spent on reports, wind tunnel model, and iron bird construction. In the spring semester $398.24 was spent on aircraft construction. This brings the total cost of the Raven project to $663.78.
After the initial first flight attempt it was apparent that additional repair costs could prove to be significant. Therefore, a second budget was started specifically for repairs. This repair budget included all items purchased after 24 March 97 for aircraft reconstruction. The reconstruction totaled $147.97. This brings the reconstruction to $811.75. A summary of this budget report can be found in table 16.11.1.
12 Concluding remarks
1 Conclusions
By the end of flight testing, we had devoted two entire semesters to the design, analysis, construction, and flight testing of an RPV, the Raven. We could then draw the following conclusions:
( Met all design criteria.
( Aircraft weighed slightly less than predicted values
( Project completed under budget
( Time constraints met. Aircraft was ready to fly on 1 April 1997, although weather delayed the actual flight until 2 April.
( By achieving sustained flight and landing the aircraft in a condition in which it could fly again, the ring-tail was proven as a feasible tail configuration.
( The wind tunnel tests predicted, and flight testing confirmed, that although the ring tail was a feasible empennage configuration, it was extremely sensitive in pitch to varying flow conditions. Hence a pitch controller would enhance the flight capabilities of the aircraft.
( The four inch chord ring-tail needs to be examined further, since analysis predicted it should have been capable of providing sufficient control power for flight but this was obviously not the case.
2 Future Research
Given the peculiar flight characteristics exhibited by both the four and six inch chord ring-tails, future work will involve wind tunnel testing of the full-size aircraft at varying angles of attack and sideslip. It is hoped that this will further explain the unique flow characteristics through the tail.
Although the ring-tail has proven to be very sensitive in pitch, it is clear that it does provide sufficient control power. However, the aircraft lacks yaw control. A possible solution is to also gimbal the tail in yaw.
13 Acknowledgments
We want to thank several people for their contributions to helping complete this project. Without their help we could have never accomplished all that we did. Thank you to:
( Dr. Perkins, our professor, advisor, and ultimate source of motivation
( Dr. Hall, our stability and controls advisor and electronics consultant
( Dr. Yuan, our ANSYS and structures advisor
( Robert Vess, our lab technician and ace pilot
( Chris Gibson, our teaching assistant
( Scott Newbern and Scott Davis, our graduate advisors
( Stearns Heinzen, our undergraduate advisor
( Todd Davis, for taking the time to make us a new ring-tail when we were about out of time
( The North Carolina Supercomputing Center, our ANSYS savior after the EOS administration let the ANSYS license expire early in the fall for the umpteenth year in a row
( The Graduate students of the NCSU Hypersonics Lab, for high-speed computing and logistical support for our final presentation slides
14 References
ASM International. Engineered Materials Handbook, Vol. 1-Composites,
Metals Park, OH, 1987.
Etkin, Bernard and Reid, Lloyd D. Dynamics of Flight: Stability and Control, 3rd ed. John Wiley & Sons, Inc., New York, NY, 1996.
NCSU MAE Department. Final Report: Raven Design Team, Vol. 1, MAE 478-Senior Design, Raleigh, NC, 1996.
Raymer, Daniel P. Aircraft Design: A Conceptual Approach, 2nd ed. American Institute of Aeronautics and Astronautics, Inc., Washington, DC, 1992.
Trayer, George. Wood in Aircraft Construction, The National Lumber Manufacturers Association. Washington, DC, 1930.
15 Appendix
1 Program Trim
***************************************************************************
* *
* This program will calculate the trim angles for the angle of attack and the *
* tail of an aircraft. Also, it will produce data for several elevator angles*
* in order to plot this data on a moment coefficient versus angle of attack *
* graph. *
*
********************************************************************************
*variables
REAL Cmo
REAL CLde
REAL Cmde
REAL CLtrim
REAL CLalpha
REAL Cmalpha
REAL alpha
REAL CL
REAL de
REAL W
REAL rho
REAL uo
REAL S
REAL pi
REAL it
REAL df
REAL CLdf
REAL Cmdf
REAL CLo
REAL CL
REAL SM
REAL alphaflap
*variable constants
pi=3.14159
rho=.002377
S=384.0/144.0
CLalpha=5.404
CLde=.24
Cmde=-1.252
Cmalpha=-.539
CLdf=2.019
Cmdf=-0.2204
Clo=.3444
Cmo=-0.03951
*asking the user for some variables
WRITE(*,*)'What is the weight of aircraft(lbs)='
READ(*,*)W
WRITE(*,*)'What is the speed of the aircraft(ft/s)='
READ(*,*)uo
WRITE(*,*)'What is the flap deflection(deg)='
READ(*,*)df
WRITE(*,*)'What is the static margin(%)='
READ(*,*)SM
CLtrim=(W/(0.5*rho*(uo**2)*S))
CL=CLtrim
SM=SM/100
df=df*pi/180.0
*calculation
it=((SM*Cmalpha/CLalpha*(CLo-CL))-Cmo)/(Cmde-(SM*Cmalpha/Clalpha*CLde))
IF(df.NE. 0.0)THEN
de=-1.0*(Cmo+Cmdf*df+Cmde*it+(SM*Cmalpha/CLalpha)*(CL-CLo-
$ CLdf*df-CLde*it))/(Cmde-(SM*Cmalpha/CLalpha*CLde))
alphaflap=(CL-CLo-CLdf*df-CLde*(de+it))/CLalpha
de=de* 180.0/pi
it=it* 180.0/pi
alphaflap=alphaflap* 180.0/pi
df=df* 180.0/pi
WRITE(*,*)'The tail incidence for',uo,'ft/sec is',it,
$ 'degrees'
WRITE(*,*)'The elevator deflection is',de,'degrees'
WRITE(*,*)'The angle of attack is',alphaflap,'for flap
$ deflection of',df,'degrees'
ELSEIF(df.EQ. 0.0)THEN
it=((SM*Cmalpha/CLalpha*(CLo-CL))-Cmo)/(Cmde-(SM*Cmalpha/Clalpha
$ *CLde))
alpha=(CL-CLo-CLde*it)/CLalpha
it=it* 180.0/pi
alpha=alpha* 180.0/pi
WRITE(*,*)'The tail incidence for',uo,'ft/sec is',it,
$ 'degrees'
WRITE(*,*)'The angle of attack is',alpha,'degrees'
ENDIF
END
16 Tables
1 Introduction
2 Design History
3 Wind Tunnel Testing
1 New Pitch Rate and Yaw Rate Coefficients
|Case |CL |Cm |Cy |Cn |Cl |
|q = 2.73 rad/s |0.3971 |-0.1748 | | | |
|q = 1.638 rad/s |0.3541 |-0.1992 | | | |
|(landing speed) | | | | | |
|r = -3.92 rad/s | | |-0.0129 |0.00703 |-0.005425 |
* Coefficients left out are insignificant and are not considered in stability analysis
4 Stability and Controls
1 Stability Derivatives for 4” Tail Chord
|Longitudinal | |Lateral | |
|Velocity = 65 mph | |Velocity = 65 mph | |
|CL( |5.404 |Cy( |-0.2826 |
|Cm( |-0.5390 |Cn( |0.08385 |
|CLq |11.91 |Cl(` |-0.02629 |
|Cmq |-30.64 |Cyp |-0.1226 |
|CL(e |0.2400 |Cnp |-0.0042 |
|Cm(e |-1.252 |Clp |-0.8094 |
|Cm(dot |-2.056 |Cyr |0.3093 |
|CLo |0.5208 |Cnr |-0.0843 |
|CDo |0.0400 |Clr |0.0650 |
|Cxu |-0.0800 |Cy(a |-0.0561 |
| | |Cn(a |0.0351 |
| | |Cl(a |-0.4959 |
| | |Cn(d |-0.0141 |
| | |Cy(d |0.0705 |
| | |Cl(d |0.0434 |
* All derivatives are per radian
2 Trim Angles for Varying Flight Configurations
| |Weight (lbs) |Speed |Flap Deflection |Static Margin |Tail Incidence |Elevator |Angle of Attack |
| | |(ft/s) |(deg) |(%) |(deg) |Deflection (deg) |(deg) |
|Takeoff with Flaps |10.6 |70 |10 |16 |-2.0613 |-1.507 |7.74 x 10-3 |
|Takeoff without Flaps |10.6 |70 |0 |16 |-2.0613 |_____ |3.677 |
|Flight with Flaps |10 |95 |10 |10 |-1.814 |-1.603 |-3.529 |
|Flight without Flaps |10 |95 |0 |10 |-1.814 |_____ |0.136 |
|Landing with Flaps |10 |70 |10 |10 |-1.949 |-1.603 |-0.403 |
5 Structures
1 Material Properties
| |Balsa |Carbon Fiber/Epoxy |
|Ex (psi) |9.1375 ( 105 |21.0 ( 106 |
|Ey |1.595 ( 104 |1.7 ( 106 |
|Ez |4.351 ( 104 | |
|Gxy (psi) |2.901 ( 104 |0.65 ( 106 |
|Gxz |4.496 ( 104 | |
|Gyz |4.786 ( 103 | |
|( (lb/in3) |7.20 ( 10-3 |5.6 ( 10-2 |
|((ten)max (psi) |3.34 ( 103 |180.0 ( 103 |
|((comp)max (psi) |1.74 ( 103 |180.0 ( 103 |
* Carbon was considered an isotropic material, and, as such, not all material properties are necessary for analysis.
2 Summary of ANSYS Results
| |Wing Tip Deflection (inches) |Maximum (y (psi) |Maximum Von Mises Stress (psi) |
|(= 0(, 1g |0.269 |-12732 |13975 |
|(= 0(, 6g |1.613 |-76392 |83847 |
|(= 5(, 1g |0.271 |-12639 |13900 |
|(= 5(, 6g |1.641 |-75834 |83436 |
|(f = 5( |0.256 |-11872 |13039 |
|(f = 30( |-0.381 |13704 |14100 |
6 Weights and Balances
1 Predicted vs. Actual Weights
|Component |Predicted Weight (lbs) |Actual Weight (lbs) |
|Engine System |2.08 |2.1 |
|Fuselage |2.7 |2.6 |
|Instrumentation |4.06 |2.91 |
|Wing |2.05 |1.45 |
|Control Surfaces |0.63 |0.56 |
|Ring-tail |0.53 |0.24 |
|Total |12.05 |9.9 |
2 Weight and Balance Summary
|Raven Dry Weight |Longitudinal Ballast required |
|Estimated Fall 1996 |12.05 lbs | |
|First Flight 1997 actual |9.9 lbs |0.1 lbs |
|Second Flight 1997 actual |9.5 lbs |0.2 lbs |
|Third Flight 1997 actual |10.1 lbs |0 lbs |
|Fourth Flight 1997 actual |10.1 lbs |0 lbs |
7 Performance
1 Parameter Values
|e = 0.7731 |S = 2.667 ft2 |
|k = 0.04392 |cLdf= 2.019 rad-1 |
|cDo = 0.03 |cLa= 5.5 rad-1 |
|cD = cDo+k*cl2 |cLmax (w/o flaps) = 1.363 |
|air density at sea level = |cLmax (w flaps) = 1.661 |
|2.3769E-3 slug/ft3 | |
|weight = 12 lb | |
2 Performance Summary
| |Takeoff |Cruise |Approach |Touchdown |Stall |
|V (ft/s) |70.7 |95.3 |76.6 |67.8 |58.9 |
|cL |0.9085 |a=0 : 0.5211 |0.8044 |1.027 | |
| | |L=W : 0.4169 | | | |
|cD |0.06625 |a=0 : 0.04193 |0.05841 |0.07632 | |
| | |L=W : 0.03763 | | | |
8 Construction
9 Flight Testing
1 Takeoff Settings for Individual Flights
|Takeoff Settings |Flaps |Tail Incidence |Dolly Incidence |Flaperon Rate |Deflector Rate |Elevator Rate |
|Flight #1 |10 ( |4( |6( |Medium |High |High |
|Flight #2 |10( |0( |4( |Medium |High |High |
|Flight #3 |10( |0( |1.5( |Medium |High |High |
|Flight #4 |10( |0( |1.5( |Medium |High |High |
10 Modifications and Repair
1 Effects of Vertical Fins
|Configuration |(Cnb)ring |(Cnb)fins |(Cnb)aircraft |
|No Fins |0.157 |N/A |0.08385 |
|Fins |0.0765 |0.047 |0.06965 |
|Coefficient |Absolute Change of Fins |Percent Change of Fins |
|(Cnb)ring |-0.0805 |-51.27% |
|(Cnb)aircraft |-0.0142 |-16.94% |
* All derivatives are per radian.
2 Effect of Servo Boxes
|Coefficient |Value without Boxes |Value with Boxes |Absolute Change of Boxes |Percent Change of Boxes |
|CL |0.3444 |0.2612 |-0.0832 |-24.16% |
|Cm |-0.03951 |-0.02869 |0.01082 |-27.39% |
|CLa |5.404 |5.295 |-0.109 |-2.02% |
|Cma |-0.539 |-0.4349 |0.1041 |-19.31% |
|CLb |-0.1 |0.125 |0.225 |-225% |
|Cyb |-0.2826 |-0.3087 |-0.0261 |9.24% |
|Cmb |0.1143 |0.0057 |-0.1086 |-95.01% |
|Cnb |0.08385 |0.08995 |0.0061 |7.27% |
|Clb |-0.026285 |-0.043945 |-0.01766 |67.19% |
* All derivatives are per radian.
3 Aerodynamic Coefficients of Final Configuration
|Case |CL |Cm |Cy |Cn |Cl |
|Baseline |0.3439 |-0.04154 |0 |0 |0 |
|a = 0.1 rad |0.8929 |-0.1343 | | | |
|b = 0.1 rad | | |-0.0328 |0.01067 |-0.002234 |
|p = -3.92 rad/s | | |0.0103 |0.001732 |0.0833 |
|q = 2.73 rad/s |0.6419 |-1.4002 | | | |
|q = 1.638 rad/s** |0.4368 |-0.305 | | | |
|r = -2.73 rad/s | | |-0.0261 |0.014 |-0.009315 |
|de = 0.1 rad |0.3774 |-0.2134 | | | |
|dd = 0.087 rad | | |-0.00665 |0.001474 |-0.003804 |
|da = -0.087 rad | | |0.0055 |-0.003369 |0.04325 |
|df = 0.087 rad |0.5201 |-0.05903 | | | |
* Coefficients left out are insignificant and are not considered in stability analysis.
** This case was run at landing speed. All other cases are for cruise speed.
4 Stability Derivatives 6” Tail Chord
|Longitudinal | |Lateral | |
|Velocity = 65 mph | |Velocity = 65 mph | |
|CL( |5.490 |Cy( |-0.3280 |
|Cm( |-0.9276 |Cn( |0.1067 |
|CLq |39.04 |Cl(` |-0.2230 |
|Cmq |-177.9 |Cyp |-0.3191 |
|CL(e |0.3350 |Cnp |-0.0084 |
|Cm(e |-1.719 |Clp |-0.4052 |
|Cm(dot |-2.870 |Cyr |0.3646 |
|CLo |0.5208 |Cnr |-0.0978 |
|CDo |0.0400 |Clr |0.0651 |
|Cxu |-0.0800 |Cy(a |-0.0630 |
| | |Cn(a |0.0193 |
| | |Cl(a |-0.2478 |
| | |Cn(d |-0.0084 |
| | |Cy(d |0.0973 |
| | |Cl(d |0.0218 |
* All derivatives are per radian
11 Budget Overview
1 Initial and Repair Budgets
|Initial Budget Items |Cost |
|Reports |$87.05 |
|Wind Tunnel Ring |$24.25 |
|Balsa Wood |$38.57 |
|Body Fillers |$53.17 |
|Adhesives |$82.77 |
|Ply Wood |$52.16 |
|Dowels |$8.67 |
|Paint (Primer) |$17.64 |
|12 oz. Fuel Tank |$3.19 |
|Dubro Servo Arms |$9.19 |
|3-APC 12x7 Props |$7.38 |
|Sullivian Push Rod |$5.59 |
|2-Robart Hinge Points |$8.18 |
|9- 40-40 Ball Links |$28.61 |
|60 Size Engine Mount |$19.95 |
|2" Tru Turn Spinner |$25.20 |
|R/C Foam Padding |$8.37 |
|2-Servo Extensions |$20.00 |
|Misc. Engine Supplies & Wood |$49.57 |
|Construction Aids |$41.76 |
|Sand Paper |$33.30 |
|Hardware |$39.21 |
| | |
|Total |$663.78 |
| | |
|Repair Budget Items | |
|2-Monokote |$23.04 |
|3- Dubro R/C Foam |$8.87 |
|Robart Hinges |$6.35 |
|2" Tru Turn Spinner |$21.19 |
|Engine Mount |$3.17 |
|Hardware |$29.69 |
|Adhesives |$23.19 |
|Balsa Wood |$5.58 |
|Props & Fuel |$26.89 |
| | |
|Total |$147.97 |
|Overall Total |$811.75 |
17 Figures
1 Int
2 Design History
1 Overall Configuration with 4” Tail
2 Overall Configuration with 6” Tail
3 Wind Tunnel Testing
1 Wind Tunnel Testing Setup
2 Aircraft Mesh with Deflectors and Resized Flaperons
3 Pressure Contour with Deflector
4 Pressure Contour with Ailerons
4 Stability and Controls
1 Aileron and Deflector Comparison for 0.1 rad Doublet
2 Block Diagram of Yaw Damper
3 Yaw Rate Comparison for Baseline and Yaw Controller
4 Block Diagram of Non-linear Servo
5 Structures
1 Fuselage Internal Structure
2 Finite Element Mesh of Wing
3 Wing Tip Deflection at ( = 0(, 1g Loading
4 Wing Tip Deflection at ( = 0(, 6g Loading
5 Wing Tip Deflection at ( = 5(, 1g Loading
6 Wing Tip Deflection at ( = 5(, 6g Loading
7 Wing Tip Deflection with (f = 5(
8 Wing Tip Deflection with (f = 30(
9 (y at ( = 0(, 1g Loading
10 (y at ( = 0(, 6g Loading
11 (y at ( = 5(, 1g Loading
12 (y at ( = 5(, 6g Loading
13 (y at (f = 5(
14 (y at (f = 30(
15 Von Mises Stress at ( = 0(, 1g Loading
16 Von Mises Stress at ( = 0(, 6g Loading
17 Von Mises Stress at ( = 5(, 1g Loading
18 Von Mises Stress at ( = 0(, 6g Loading
19 Von Mises Stress at (f = 5(
20 Von Mises Stress at at (f = 30(
6 Weights and Balances
1 Internal Component Layout
7
8 Construction
1 Sample Fuselage Cross-Section
2 Completed Templates and Cross-Section Layout
3 Templates Mounted on Blue Foam
4 Joined Plug
5 Top Half of Mold
6 Original Single Spar Wing Cross-Section
7 Wing Layout
8 Wing Servo Fairings
9 Wing Mounting System
10 Ring-Tail Foam Core Template
11 Ring-Tail Facet Cross-Section
12 Ring-Tail Mounting Mechanism
13 Ring-Tail Facets in Assembly Template
14 Paint Scheme
15 Fuselage After Coat of White Paint
16 Wing After Coat of White Paint
17 Completed Raven Aircraft
9 Flight Testing
1 Launch Dolly
2 Aircraft Prepared for Takeoff
3 First Flight Attempt Immediately After Takeoff
4 Wing Damage Sustained After First Flight Attempt
5 Ring-Tail Damage Sustained After First Flight Attempt
6 Fuselage Damage Sustained After First Flight Attempt
10 Modifications and Repair
1 Exploded View of Wing
2 Pressure Contour of Sideslip with Fins
3 Aircraft Mesh with Servo Fairings
4 Baseline Pressure Contour with Servo Fairings
5 Baseline Pressure Contour with 6” Tail
6 Ground Effect Survey
7 Pitch Angle Comparison for 0.1 rad Doublet
8 Pitch Rate Comparison for 0.1 rad Doublet
9 Yaw Angle Comparison of Ailerons for 0.1 rad Doublet
10 Yaw Angle Comparison of Deflectors for 0.1 rad Doublet
11 Roll Angle Comparison of Ailerons for 0.1 rad Doublet
12 Roll Angle Comparison of Deflectors for 0.1 rad Doublet
13 Wing Pivot Mechanism
18 Index
A
Adverse yaw 10
Aerodynamic coefficients 10, 17, 32, 61
Aerodynamics 1, 10, 15, 17, 28, 31, 32, 61
Ailerons 10, 11, 32, 73, 75, 135, 137
AutoCAD 1
B
Balsa 2, 13, 14, 22, 23, 24, 25, 31, 48, 64
Birch 24
Bondo 19, 21, 22
Budget 2, 34, 35, 63, 64
Bulkhead 20, 21
Bungee 17, 27
C
Carbon 8, 13, 14, 15, 20, 22, 24, 33, 48
Chord 8, 13, 22, 23, 24, 25, 28, 29, 32, 35, 45, 62
Compression 14, 22
Control horn 26
Control surfaces 8, 10, 27
Controller 11, 35, 77
Cost 34, 64
D
Deflectors 8, 10, 11, 28, 29, 32, 57, 71, 72, 75, 136, 138
Design loads 7, 9
Digital cards 16, 34
Dolly 7, 22, 27, 28, 29, 57, 120
Doublet 75, 133, 134, 135, 136, 137, 138
Dowel 19, 24, 64
Drag 8, 10, 11, 17, 24
Drag model 17
Drag rudders 8, 10, 11
E
Engine 7, 21, 27, 28, 33, 34, 51, 64
Engine mount 64
Epoxy 20, 21, 22, 23, 24, 34, 48
F
Fairing 8, 9, 24, 25, 28, 31, 32, 109, 129, 130
FDR 16, 34
Fiberglass 15, 16, 19, 20, 21, 23, 25, 31, 32
Flaperons 8, 10, 23, 28, 32, 57, 71
Flaps 8, 10, 14, 15, 17, 39, 46, 54, 57
Flight test 2, 7, 8, 9, 11, 13, 15, 26, 27, 33, 35
Flocking 21, 22, 24
Foam 2, 15, 18, 19, 20, 21, 22, 23, 24, 25, 31, 64, 104, 111
Formica 18, 20, 21, 22
Fuel 64
G
Gust 30
H
Hatches 20, 21, 22, 25
Hot wire 18, 19, 22, 23
I
Internal systems 21
IRON BIRD 24, 34
K
Kevlar 13, 20, 21
L
Landing 7, 9, 15, 17, 29, 30, 35, 43, 46, 61
Linear 11, 12, 78
Linkage 24
Load 7, 8, 9, 13, 14, 15, 23, 28, 29, 33
M
Mass balance 23
Material properties 13, 48
Meshes 13, 32, 71, 81, 129
Microballoons 21, 22
Mold 18, 19, 20, 21, 22, 106
Moment of inertia 16
P
Performance 1, 17, 18, 27, 28, 53, 55
Pitch 7, 10, 29, 30, 32, 35, 43, 133, 134
Plug 18, 19, 105
Plywood 21, 22, 25
PMARC 8, 10, 11, 28, 32
Poplar 13, 21, 22, 23
Procurement 1
Propeller 7, 33
R
Range 2, 27
Resin 16, 23, 24, 34
Roll 8, 10, 11, 29, 32, 137, 138
S
Safety factor 7
section 10, 13, 14, 18, 19, 20, 22, 24, 25
Separation 32
Servos 8, 9, 11, 22, 23, 24, 28, 31, 33, 34, 60, 64, 78, 109
Sheeting 2, 23, 24, 25
Skids 7, 9, 16, 21, 22, 29, 30, 33
Skin 13, 18, 20, 21, 22, 28
Spars 2, 8, 13, 14, 15, 22, 23, 24, 28, 31, 33, 107
Spinner 64
Stability Augmentation System (SAS) 16, 21, 34
Stall 55
Static margin 7, 16, 39
Strain 14
Stress 13, 14, 15, 33, 49, 94, 95, 96, 97, 98, 99
Stringers 13, 20
T
Tail boom 22, 25
Takeoff 17, 27, 28, 29, 46, 55, 57, 121, 122
Template 18, 24, 25, 104
Tension 14
Trim 12, 22, 38, 46
U
Unigraphics 1, 26
V
Velocity 11, 17, 45, 62
Von Mises Stress 49, 94, 95, 96, 97, 98, 99
Vortex 10
W
Weight 1, 7, 13, 16, 17, 22, 23, 39, 46, 50, 51, 52, 54, 100
Wing loading 8
Wing mounting system 14
Wing tip 9, 14, 22
Wing tips 22
Y
Yaw 2, 8, 10, 11, 28, 29, 31, 32, 35, 43, 76, 77, 135
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