Complete Design Review
Project 5008
Christina Alzona
Benjamin Wagner (team leader)
2/18/05
1.0 Introduction
1.1 Design Development
1.2 Design Alternatives
2.0 Conceptual Design
2.1 Needs Assessment
2.2 Feasibility Assessment
2.2.1 Aircraft Type
2.2.2 Empennage
2.2.3 Landing Gear
2.2.4 Propulsion System
2.2.5 Wings
2.2.6 Radio and Transmitter
2.2.7 Building Materials and Construction Methods
2.3 Conclusions
3.0 Preliminary Design
3.1 Construction Methods
3.1.1 Wing
3.1.2 Empennage
3.1.3 Fuselage
3.1.4 Takeoff Mechanism
3.2 Analysis Methods and Sizing
3.2.1 Aerodynamics
3.2.2 Structures
3.2.3 Payload
3.2.4 Propulsion
4.0 Prototype Design
4.1 Construction Methods
4.1.1 Wing
4.1.2 Empennage
4.1.3 Fuselage
4.2 Aircraft Configuration
4.2.1 Propulsion System
4.2.2 Weight Analysis
4.3 Predicted Performance
5.0 Analysis and Design Testing
5.1 Wing Failure Analysis
5.2 Aircraft Load Testing
5.3 Center of Gravity Calculations
6.0 Bill of Materials
7.0 Time Line
7.1 Fall 2004 Timeline
7.2 Winter 2004 Timeline
7.2.1 Predicted Winter 2004 Timeline
7.2.2 Actual Winter 2004 Timeline
8.0 Acknowledgements
9.0 References
Appendix A: Construction Pictures
Appendix B: CAD Drawings
Appendix C: Finite Element Analysis
Appendix D: Electric Motor Calculations
Appendix E: Airfoil Data
1.0 Introduction
Rochester Institute of Technology College of Imaging Arts and Sciences had expressed an interest in the creation of an unmanned airborne sensing platform to assist their Wildfire Airborne Sensor Program (WASP). In an effort to assist CIAS, two design teams were assembled. One will design and build the body of the UAV, and another team will design and integrate onboard telemetry and stability augmentation. This UAV approach is scheduled to be an ongoing project in coming years. Current mission objectives include flight ranges of at least two miles from base station and an endurance of at least one hour. CIAS also requested a fairly large payload capacity and ease of repeatability in design and construction. CIAS has requested a UAV capable of a 3.2 km range, able to carry a 1.5 kg payload, and adaptable to various unspecified mission requirements.
1.1 Design Development
During the conceptual design phase, aircraft size approximations and feasibility assessment were developed from the needs of our sponsors in the Mechanical Engineering department and CIAS as RIT. Using the initial size approximations, feasibility assessment and the needs of our sponsors, the team prioritized the most critical elements of the project for further consideration. These elements were then analyzed based on aerodynamics, propulsion and structural integrity of the aircraft. These categories were further divided into building materials, aerodynamic designs, structural designs, and construction methods. In this phase, numerous airplane configurations were considered using our design parameters and compared against our feasibility assessment of the different aspects of the aircraft design. The models that most closely achieved the desired design criteria were later utilized in the preliminary design phase for closer examination.
In the preliminary design phase, a more detailed analysis and theoretical performance calculations were performed. Flight characteristics were predicted and loads were analyzed for the optimal design. Payload volumes and weights were dynamic throughout the preliminary design phase making the fuselage difficult to conceptualize, but the final fuselage design should meet the specifications that were given as of this preliminary design review. Construction methods for the entire airborne platform had to take into account the limited human labor that will be available with the current team. It is the recommendation of the team that more members be allocated to the construction of the airborne platform to ensure the completion of the prototype in the time constraints imposed.
1.2 Design Alternatives
Numerous propulsion, aerodynamic, and structural configurations were compared and analyzed. Due to existing molds and past experience the airfoil for the main wing was chosen to be an Eppler 423. Main wing locations that were considered were high, low, and mid fuselage mounts. A high wing location was chosen because of increased stability over other the designs and it would allow the integrating teams to access hardware more easily within the fuselage. Empennage styles that were considered were conventional, T-tail, cruciform, and canard. The conventional tail was chosen for ease of constructability and to keep weight down. Propulsion options that were analyzed were electric, hybrid, and nitromethane glow engine. Electric was chosen to meet the vibrations specifications of the needs assessment, it produces no emissions to affect the integrity of the video equipment onboard, and it increases stability because as fuel is consumed the weight of the airborne platform does not shift. Motor location was chosen to be a front mount. The front mounted position allowed more air to pass over the wing to aid lift during take-off and kept weight down as other options would have required reinforcing the airframe. Lithium polymer batteries were chosen as a fuel source because currently it offers the highest energy density per unit of mass of any commercially available batteries. The main drawback of lithium polymer is the cost. A fiberglass and epoxy composite was chosen for shell of the fuselage. It provides the strength necessary for the chosen geometric configuration but is less stiff than other options to better achieve the vibrations requirement.
2.0 Conceptual Design
The conceptual design process was separated into two main tasks. The first task was to determine the needs assessment to evaluate the minimum requirements of our sponsor and define the goals of the design team. Secondly, a feasibility assessment was performed on all the main aspects of the airborne platform design. With the results of the feasibility assessment, a preliminary design incorporating the winning designs from each aspect was created.
2.1 Needs Assessment
This design partnership needed to take into consideration that it would be integrating the airborne platform design with at least two other design teams in the winter and spring quarters. These two other projects that will be involved in adding sensitive and expensive equipment of a volume and weight that was yet to be determined as of this preliminary design review. The design of the airborne platform had to incorporate these two black boxes comfortably and allow their components to be accessed easily. The airborne platform also had to provide some measure of impact protection of its more expensive components in the unfortunate event of a crash.
In addition to the above concerns, CIAS has specified that the airborne sensing platform must meet or exceed these specifications:
· Carries a 3 lb payload (CIAS sensing equipment)
· Has a cruise speed of 15-30MPH
· Has a 1 hour endurance
· Provides view angles both upward and downward for sensors
· Can climb to 1,000 ft
2.2 Feasibility Assessment
With the needs assessment completed. The team split the design into seven main categories to brainstorm ideas. Once the team exhausted itself of the best options for each category, a feasibility assessment was used to rate them against one another. The best design from each category was then incorporated into the preliminary design.
2.2.1 Aircraft Type
Several aircraft configurations were analyzed during the conceptual design phase of this project. These configurations were conventional, flying wing, canard, and biplane as illustrated in figure 1.
Figure 1 Aircraft Type
Aircraft configurations were chosen based on general aircraft knowledge and experience. The conventional configuration was chosen because of stability and its ability to contain the unknown volumes of the payloads being carried. Although a flying wing design may have been more efficient to reduce drag, the unknown nature of the payload made it difficult to design a volume to contain it. The canard design was dismissed because the downwash of the horizontal stabilizer of this design has been proven to disrupt the lift distribution on the wing, thereby increasing the induced drag and shed vorticity.
2.2.2 Empennage
Several empennage configurations were considered for the design of this airborne platform. A conventional style tail design was chosen because of ease of constructability and weight concerns. A T-tail and cruciform were considered because they would keep the horizontal stabilizer out of the downwash of the wing, but it was thought the airborne platform would be moving too slow for this advantage to be noticed considerably. The T-tail and cruciform designs would also require reinforcement of the vertical stabilizer and increased the weight of the airplane. The V-tail would decrease weight, but it was discounted because it would require using a radio transmitter capable of mixing control surface functions. V-tails also produce a counteracting lift which would have decreased the lift distribution of the airborne platform. A visual representation of the empennage choices are found in Figure 2 and the feasibility analysis is found in Figure 3.
Figure 2 Empennage Configurations
Tail Design Feasibility Assessment(On a scale from 1-10) / Weighting / Low / T tail / Cruciform
R1: Sufficient Skills / 0.1 / 9 / 7 / 6
R2: Sufficient Equipment / 0.1 / 9 / 9 / 9
R3: Sufficient # of people / 0.13 / 2 / 2 / 2
E1: Economically Feasible / 0.07 / 8 / 7 / 7
S1: Meeting Intermediate Milestones / 0.1 / 8 / 8 / 8
S2: Meeting PDR Requirements / 0.1 / 8 / 7 / 7
S3: Meeting CDR Requirements / 0.15 / 8 / 7 / 6
T1: Has similar technology been used before / 0.12 / 8 / 7 / 6
T2: Plane stability / 0.08 / 6 / 8 / 7
T3: Drag reducing / 0.05 / 7 / 7 / 7
Total: / 1 / 7.21 / 6.73 / 6.28
Figure 3: Tail design feasibility
2.2.3 Landing Gear
The landing gear configurations that were considered were tail-dragger, tricycle gear, and no landing gear. A tail-dragger configuration was initially considered optimal for this airborne platform because it provided an decent trade-off between aerodynamic drag and ground stability. While a tricycle landing gear would have been the most stable in ground handling characteristics, the large cross section of the nose gear would have added and unacceptable amount of drag to the flight characteristic predictions. No landing gear was also considered and was eventually chosen. By choosing no landing gear, the design would be save weight and be more aerodynamic. Onboard fuel would also be conserved because an external source of power would have to be used to launch the aircraft. A method of skids or other devices would be need, however, to protect the fuselage and other components against a rough belly landing. A diagram of the feasibility analysis is found in Figure 4.
Landing Gear(On a scale from 1-10) / Weighting / None / Trike / Tail-dragger
Aerodynamics / 0.2 / 9 / 7 / 8
Ground Handling / 0.2 / 0 / 9 / 6
Weight / 0.3 / 10 / 6 / 8
Ease of Construction / 0.3 / 10 / 6 / 7
Total: / 1 / 7.8 / 6.8 / 7.3
Figure 4: Landing gear feasibility analysis
2.2.4 Propulsion System
Several propulsion systems were examined. These included nitromethane glow engines, electric motors, and a hybrid of the previous two options. An all electric configuration was chosen because it would produce the least vibrations of all options. It is also the cleanest and does not produce any exhaust that may affect the onboard sensory payload. The main disadvantages of electric power are the added weight of batteries and electric motors generally have a lower energy per unit mass than the glow engines. Glow engines have an advantage of a outputting more power per weight, but create an unacceptable level of vibrations. Glow engines also create an oily exhaust that may stick to the aircraft and could affect the sensing ability of the payload. As fuel is consumed with a glow engine, the center of gravity of the aircraft will shift changing the stability of the aircraft in mid-flight. A hybrid would combine some advantages of the previous options. The power available advantage of the glow engine would have been utilized in takeoff, climbing, and getting the airborne platform to the loiter zone of the flight mission. In the loiter zone and the return flight to the landing zone, the electric motor would be used to decrease the vibration level for the CIAS sensory equipment. The hybrid option would require reinforcements throughout the fuselage to support both power plants, would still create exhaust that could interfere with sensory equipment, and would have the stability issues associated with mid-flight center of gravity shifts. Analysis of the propulsion feasibility can be found in figure 5.
Propulsion(On a scale from 1-10) / Weighting / Electric / Hybrid / Gas
Vibration / 0.3 / 9 / 7 / 5
Weight / 0.2 / 8 / 9 / 7
Stability / 0.2 / 9 / 8 / 7
Cost / 0.1 / 7 / 7 / 7
Ease of Installation / 0.2 / 9 / 7 / 8
Total: / 1 / 8.6 / 7.6 / 6.6
Figure 5: Propulsion feasibility analysis
An issue that developed when an all electric configuration was chosen was the option of battery type. After researching what was commercially available at the time of this preliminary design review, the options that were available were nickel metal hydride, nickel cadmium, and lithium polymer. Lithium polymer was chosen because its energy density is much higher than the other two options. Also, the discharge curve of lithium polymer, voltage over time, is largely flat until the battery is completely discharged. Battery weight and volume were the main concerns considering the largely unknown black box payload. The main disadvantages of lithium polymer batteries are the cost and they have a slower charge rate than other options. Nickel metal hydride and nickel cadmium were much cheaper but weight much more than the lithium polymer option. The voltage supplied by these batteries also steadily declined during the discharge cycle. An analysis of the battery feasibility is shown in figure 6.
Batteries(On a scale from 1-10) / Weighting / LiPoly / Nimh / NiCad
Energy Density (charge/mass) / 0.4 / 9 / 7 / 8
Cost / 0.1 / 6 / 8 / 7
Ease of use/installation / 0.2 / 8 / 8 / 8
Recharge Ratio / 0.3 / 7 / 7 / 7
Total: / 1 / 7.9 / 7.3 / 7.6
Figure 6: Battery feasibility analysis