Conceptual Design Requirements of a Newly Developed UAV as a Research Platform at KFUPM: Structure, Aerodynamics and Propulsion Amro M. Al-Qutub* King Fahd University of Petroleum and Minerals, Dhahran 31261, Saudi Arabia
The objective of the present paper is to set the structure, aerodynamic and propulsion conceptual design requirements of a multi-role long endurance UAV. The aircraft is designed to be used as a research platform at King Fahd University of Petroleum and Minerals, Mechanical engineering department. The maximum payload and takeoff weight were set at 10 kg and 65 kg, respectively, which are the main design constrains of the system. Further specifications include cruise speed of 220 km/hr, 8 hours endurance and a ceiling of 18,000 ft. This would allow the UAV to perform over 1,700 km of non-stop flight with maximum payload and without refueling. The present study was based on available information in the literature and published specifications of different UAV systems manufacturers. Results included systems weight distribution, and limitations on requirements of the aerodynamic performance and propulsion system. Design sensitivity analysis resulted in the definition of basic performance limitations to achieve the objective of the present UAV. This included high aerodynamic performance (L/D > 13) which requires a retractable landing gear configuration. Also, the propulsion system should be equipped with a pusher propeller driven by a 2 stroke IC engine (17 -22 hp) with specific power higher than 2.5 hp/kg and low SFC of 0.33 kg/hp.hr. System reliability is improved through the selection of higher engine design output. Also additional heat-sink system for engine head would further improve the service life during hot local season. The overall propulsion system performance will be enhanced with a high efficiency propeller (> 50%) during cruise. The structure should be made of light composite material to comply with limited aircraft dry weight, of 46 kg. Wings are separable to allow for testing different configurations and for the ease of store and transportation. The survivability and reliability of data and system recovery is improved by a recovery system composed of a parachute and an airbag in case of emergencies. *
Associate Professor, Mechanical Engineering Department,
[email protected]. 1 SSAS UAV Scientific Meeting & Exhibition, Jeddah, Saudi Arabia (June 6, 2006) Paper No. SSAS-2006-051
I.
Introduction
T
HE Unmanned Aerial Vehicle (UAV) is a remotely piloted or self-piloted aircraft that can carry cameras, sensors, communications equipment or other payloads. UAV technology and usage actually extends beyond Military and defense market. Many Universities around the world have their own UAV development program such as University of Sydney, Clark University, Cranfield University, Georgia Tech. University1, Berkeley University, Simon Fraser University2, and others. For Example University of Sydney UAV fleet3 is composed of five different flying UAV’s: KCEXP-series, Ariel (Figure 1), Brumby, T-Wing, and Bidule miniature Air vehicle. This besides all other UAV still under development such as The Solar powered High Altitude Endurance (HALE) meteorological UAV. The University of Sydney enjoys the funding from different organizations like Booing, and BAE.
Figure1: Sydney University UAV examples: the Ariel (left) and the Brumby (right). King Fahd University of Petroleum and Minerals (KFUPM) is encouraging multidisciplinary projects where different department contribute and integrate a system in a specific technology field. KFUPM encouraged faculty to work towards a vision that may contribute to the well fear of the surrounding society and the enhancement of the dissemination of knowledge. The development of UAV in KFUPM has the following mission: 1. Develop know-how of UAV for local applications 2. Provide experimental platform for multi-disciplinary research related to UAV applications 3. Provide a tool for academic excellence in different areas. The specific UAV should satisfy the following: • able to accommodate for equipments related to multi-disciplinary research and development work. • safe to operate • reliable for local environment • comply with local regulation of communication Building a UAV demands a vast knowledge of base vehicle platforms, flight dynamics, control theory, and real-time software in a network environment. The flight system enables the aircraft platform to serve as the autonomous agent. It requires sophisticated navigation
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sensors and integration methodologies for accurate and reliable autonomous operation. Key UAV system technologies were identified to be in: airframes; propulsion units; autonomous flight controllers; launch and recovery; navigation and guidance; selfprotection; ground control stations; payloads; and data communication, storage, processing, and dissemination (Information Technology). Key system technologies cover the following items: • UAV airframe, structure and aerodynamics • Propulsion unit • Control system (Flight control, Propulsion control) • Telemetry and ground station • Instrumentation for performance evaluation and control • Support system The objective of the present paper is to study the specific conceptual design and performance requirements of the UAV structure, aerodynamics and propulsion system.
II.
General Requirements
1)
Structure and Aerodynamics general requirements:
2)
• Composite structure + Aluminum alloy • Payload of 10 kg with gross takeoff weight of 65 kg • Reinforced • Landing gears retractable • Support up to 5g’s • Internal fuel tank and equipments • Conventional high wing configuration for stability Flight general requirements:
3)
• Take off and landing on paved ground • 160-220 Km/h cruise speed • 8+ hours endurance • Ceiling 6 km (18000 ft). Propulsion general requirements:
• Internal combustion engine (IC) high performance • Propeller driven with minimum diameter and high efficiency during cruise • If possible electrical starter + alternator • Withstand up to 45 oC, and 55 oC for short duration during takeoff • High engine efficiency with low SFC • High power/weight ratio In the design phase both engine efficiency and power/weight ratio have to be optimized for minimum system weight. Basically the payload is the most important part in the design phase since it sets the limitations on possible missions the UAV can perform. For basic research and surveillance purposes 10 kg payload is considered sufficient since continuous developments in 3 SSAS UAV Scientific Meeting & Exhibition, Jeddah, Saudi Arabia (June 6, 2006) Paper No. SSAS-2006-051
electronics avails higher performance and lower weights sensors, cameras and processors. As shown in Figure 2, compared to conventional UAV's, the payload-to-gross weight ratio
% Payload/gross Weight
PAYLOAD vs Endurance 70
Present UAV
60 50 40 30 20 10 0 3
4
5
6
7
8
9
10
Endurance (h)
Figure 2: Payload vs Endurance of available Known Conventional UAV’s ( 4) Service ceiling of UAV's
Present UAV
80,000
Ceiling (feet)
70,000 60,000 50,000 40,000 30,000 20,000 10,000 0 0
10
20
30
40
50
60
Endurance (h)
Figure 3: Ceiling vs. Endurance of available Known Conventional UAV’s4 is average, however the endurance places the present UAV among the top of the group 4. The ceiling is among the majority of UAV’s of its endurance category, Figure 3.
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A. Aerodynamics and Structure Based on the objective of the project the UAV is a long endurance (8 hrs +) with reasonable cruise speed of about (220 km/hr). Moreover, takeoff should be possible either from a launcher or a runway. This imposes a grate challenge in designing the aerodynamics and the structure of the aircraft. Endurance requires: − high aerodynamic performance (high L/D, lift to drag ratio) − Relatively Large fuel tank ( to improve weight ratio ) − Strong structure with high strength to weight ratio materials. − Low thrust specific fuel (high efficiency propulsion system) Endurance To estimate the endurance of an aircraft we may start with the mass reduction of fuel during level steady flight: dm D = −TSFC ∗ (m • g ) dt L
(1)
Where TSFC is the thrust specific fuel consumption of the propulsion system, m is the mass of the vehicle, and g is the gravitational acceleration. Assuming constant L/D and TSFC through the flight, the endurance of the vehicle can be calculated as: ∆t =
1L 1 g D TSFC
ln
mo mb
(2)
Where mo is the takeoff weight and mb is the fuel empty weight of the UAV. Takeoff requirements The UAV has to be able to takeoff from a runway, or from a launcher. Taking off from a runway requires landing gear system that reduces L/D if extended during flight. So, retractable landing gear should be used to maintain high aerodynamic performance. The major penalty in this case is increased structural weight (by only few percent). A reasonable takeoff velocity of the UAV is about 25 m/s. Based on the objective of the project, the maximum payload is 10kg. Referring to figure 3, the aircraft takeoff gross weight can be chosen as 65 kg, which corresponds to payload to gross weight ratio of 15.4%. To calculate the wing platform area, we may use the lift equation:
L=
ρ 2
V
2
CL S
(3)
With plain flap system and large aspect ratio (>9) of the wing, maximum lift coefficient can reach a value over 1.4. Using equation (3), and considering the takeoff speed to be grater than stall speed by 20%, the required wing area for the UAV for takeoff 5 SSAS UAV Scientific Meeting & Exhibition, Jeddah, Saudi Arabia (June 6, 2006) Paper No. SSAS-2006-051
will be 1.66 m2 at standard atmosphere. Higher atmospheric temperature and/or altitude would require slightly higher takeoff speed due to reduced density. For example at 50 Co, sea level, the takeoff speed will be 26.5 m/s, only a 6% increase compared to standard atmosphere conditions. Structure The Present UAV has a conventional aerodynamic configuration with pusher-type propeller See figure 4. The main fuselage has a maximum diameter of about 350 mm to contain cameras similar to the BAI5 as well as large fuel tank and Payload:
Fuselage
Propeller
Camer a Figure 4: Basic configuration of the UAV. The UAV is to be made from composites (glass fiber and/or Carbon fiber) and Aluminum alloy, to provide high strength to weight ratio. Skin parts and some of the skeleton supposed to be made from composite material while hard points and joints to be mainly from aluminum alloy for improved reliability. Part of the wing is supposed to be separable during storage and transportation for convenience of use and handling. For the systems contained in the fuselage they have to abide by the dimensional restrictions and maintain a reasonable static margin for stability purposes. Systems layout in the fuselage is illustrated in figure 5. Engine, landing gear and camera are supposed to have fixed locations. Fuel will be stored in multiple tanks, close to center of gravity for stability during flight and flexibility of payload as well as static margin consideration. The payload bay should contain flexible mount system for different parts and configuration of communication, control and payloads. The camera is (which is a part of the payload) mounted at the bottom of the aircraft as shown in figure 5. Recovery Parachute and airbag has to be on the center of gravity. 6 SSAS UAV Scientific Meeting & Exhibition, Jeddah, Saudi Arabia (June 6, 2006) Paper No. SSAS-2006-051
Fuel Tanks
Parachute
Payload bay
Airbag Propulsio n
Camera Main Landing Gear
Nose Gear
Figure 5 : Layout for Main systems in the Fuselage
Cruise performance The UAV is expected to operate at a ceiling altitude of 18,000 ft. At the cruise speed of 220 km/hr the lift coefficient is CL = 0.31 (at gross weight). This is a reasonable lift coefficient for high aerodynamic performance. Lower speed will produce higher endurance and this can be optimized according to the type of the UAV mission. Moreover, the cruise speed of the present UAV is considered among the highest in its category as can be seen in table 1 6,7. Table 1: Different UAV performance and power requirements6,7. UAV Gross Ceiling Endurance Cruise Power W/P ratio Weight (ft) (hrs) speed (hp) (kg/hp) (kg) (km/hr) Huntair 127 17,000 7.5 184 38 3.3 Raptor 854 65,000 8 50 17.1 Prowler 91 21,000 6 100 38 2.4 STM-5B 114 16,000 6 144 25.5 4.47 Pectre II 100 23,000 6 220 26 3.85 Skyeye 355 18,000 10 126 46 7.72 Hunter 727 15,000 12 162 120 6.06 Model 410 818 30,000 12 184 160 5.11 Shadow 272 17,000 14 129 52 5.23
From Table 2 the weight to power ratio of UAV's of similar category as the proposed is in the range of 2.4 to 5.2 kg/hp. The variation of designed hp on propulsion system depends on mainly on:
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− Weight of the aircraft − Flight altitude − Design clime rate − Flight speed In the present work it is proposed to use an engine with 17 to 22 hp, which corresponds to a weight to power ratio of 3.82 to 3. This will put proposed UAV among the top of the UAV's list in terms of engine power and performance. Actually the main reasons of that are the high cruise speed and the high operating temperature in the local environment, which is part of the objective for this project. To calculate the endurance of the UAV the weight ratio has to be known. Table 2 is a list of most recent UAV engines specification available in the world market that can be used for the present project8-11. Based on present technology, the propulsion system is expected to weigh about 9 kg, including engine, propeller and generator. Table.2: Some UAV Engines available in the international market. Engine Output Weight Power /Weight ratio (Brand) (hp) (kg) (hp/kg) A200B 17 5.1 / 6 kg (est.) 2.83 Quadra 300W Gen. 3W-200i B2TSQS 20 4.8/6 kg (est.) 3.67 3W Motoren W.Gen 313 22 8 2.75 Zanzotterra W. 300w Gen. L275 E 25 7.7 / 9 kg (est.) 2.78 Limbach W. Gen AR741 37 10.7 /12kg (est.) 3.1 UAV Engine Ltd. (rotary) W. Gen.
Based on the calculated size of wing and fuselage, the airframe of the UAV is expected to have a weight between 10 and 16 kg, depending on design and material used for the construction of the system. Table 3 provides the design weight of different systems in the UAV. Fuel consumption can be estimated using figure 6, which is an example for typical turboprop engine12. From this figure we can use a conservative thrust specific fuel consumption of 0.06 kg/N.hr. Using equation 2 with the assumption of having an aerodynamic performance of L/D=15, the UAV endurance will be 11.14 hrs. This satisfies the objective of the project. Table 3: Weights of different system in the UAV. Airframe
Propulsion
Control & communication
Landing Gear
Recovery System
Payload (Max.)
12 kg
9 kg
3 kg
4 kg
4 kg
10 kg
Fuel with max payload 23 kg
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Gross Weight
65 kg
Figure 6: Typical Engines TSFC (12)
The calculated endurance is based on several assumptions regarding structural technology, aerodynamic technology and propulsion system performance. Figures 7 to 9 Are the results of sensitivity analysis to illustrate the effect of each of the mentioned parameter on the endurance of the UAV. The total structure mass includes airframe, propulsion, communication and control, landing gear, and recovery system. Reducing the structure weight to 26 kg will result an increased endurance of over 15 hrs. If all typical values of aerodynamic performance and propulsion performance are met, the total structural mass has to be limited to 36 kg in order to meet the 8 hrs flight endurance Figure 7. Aerodynamic performance also proved to be critical for endurance. Increasing aerodynamic performance will increase the endurance. The limit to meet the objective here is L/D ≥13, if all other variables are typical. Reducing the (TSFC) improves the flight endurance linearly, Figure 9. Improving the TSFC down to 0.045 will increase the endurance to 14.85 hrs. Based on available technology, the TSFC has to be maintained below 0.06 during cruise conditions. Figure 10 illustrates the effect of different payload mass on flight endurance. Reducing payload to 5 kg (half of max.) will increase flight endurance to 12.3 hrs, a 1.2 hrs increase over full payload (10kg). Detailed discussion on the propulsion system is at the next section. In conclusion, there are three main parameters that affect the flight endurance: − Structure design and material − Aerodynamic design − Propulsion system performance Structure and aerodynamics are to be optimized in the design process. Propulsion performance is considered as a major design constrain.
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In conclusion the following are the aerodynamic and structure specifications: Aerodynamic: − High aerodynamic performance L/D>13 − Pay load = 10 kg max − Gross weight 65 kg − Conventional configuration − Takeoff speed = 25 m/s (standard atmosphere) − Wing Area 1.6 m2 approx. − Max cruise 220 km/hr − Endurance 8 - 12 hrs. − Max ceiling 18,000 ft − Flight range > 1,800 km
Flight Endurance ( hrs )
Structure: − Made of composites (fiberglass-carbon) with Aluminum Alloy. − Separable wings. − Retractable landing gear − Recovery system (parachute + airbag) − Total Structure weight < 36 kg − Multiple Fuel tanks − Payload rack for flexibility of mounting with min dim. of (200X200X400) mm.
16 14 12 10 8 6 4 2 0 26
29
32
35
38
41
Structure mass ( kg ) Figure 7: Effect of Structure Mass on Flight Endurance
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Flight Endurance ( hrs )
18 16 14 12 10 8 6 11
13
15
17
19
21
Aerodynamic Performance (L/D)
Flight Endurance ( hrs )
Figure 8: Effect of Aerodynamic Performance on Flight Endurance
18 16 14 12 10 8 6 0.045
0.05
0.055
0.06
0.065
0.07
0.075
TSFC ( kg/ N.hr ) Figure 9: Effect of thrust Specific Fuel Consumption on Flight Endurance
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Flight Endurance ( hrs )
15 14.5 14 13.5 13 12.5 12 11.5 11 10.5 10 10
8
6
4
2
1
Payload Mass ( kg ) Figure 10: Effect of Payload Mass on Flight Endurance, Fixed Fuel mass of 23 kg. Propulsion system The objective of the project requires a reliable and efficient propulsion system specifically during cruise condition (180 - 220 km/hr) with a ceiling of 18,000 ft. In addition, the UAV has to be practical and easy to maintain.
Type of System For the UAV application similar to present UAV category, propulsion systems are usually limited to either turbojet or propeller driven by an IC engine. Propeller UAV are more popular due to lower cost and higher efficiency at relatively low speeds (Ma20 hrs. Tolerate high g forces during catapult launch.
Cruise At level flight with a speed of V, the required propulsion power can be calculated as:
D Power P = m g V L
(4)
For cruise at 18,000 ft and 220 km/hr, with gross weight of 65 kg and minimum L/D of 11, the required propulsion power will be Powerp = 3.54 kW ( 4.74 hp ). Typically, IC engines lose about 7% of output power per 3000 ft altitude. Taking this into consideration and assuming that the engine will operate at 67% of rated power during cruise, also assuming propeller efficiency of 60%, the required engine shaft power at standard atmospheric conditions should be about 20 hp, to meet the cruise condition. The extra 33% of engine power is for maneuver and climb requirement. High altitude takeoff For takeoff at 9,000 ft the power available will be 79% of the rated engine power. To maintain a reasonable weight to power ratio of less than 5.2 kg/hp, the engine rated shaft power should be > 15.82 hp. This is even less than the engine power requirement for cruise condition. Propeller Efficiency Based on endurance calculations, the required TSFC < 0.06 kg/N.hr. Most UAV engine suppliers claim specific fuel consumption (SFC) for their product as less than 0.33 kg/hp.hr. Diesel engines recently developed for aviation application can reach SFC as low as 0.19 13. The TSFC is affected by both SFC and propeller efficiency (ηp) :
TSFC = 0.00134
V
ηP
SFC
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(5)
Propeller Efficiency (%)
Typical propeller efficiency can be as high as 85%, Figure 11. Usually each propeller design is based on a given flight speed. Designers should select propeller according to specific design requirement: - Power matching with engine - efficiency at cruise - static and low speed thrust - noise - diameter
80
60
40
20
100
200 300 400 Flight Speed MPH
500
TSFC (kg/N .hr)
Figure 11: Typical Propeller Efficiency (12)
0.08 0.07 0.06 0.05 0.04 0.03 0.02 0.01 0 0.4
0.5
0.6
0.7
0.8
Propulsion Efficiency Figure 12: Effect of Propulsion Efficiency on Thrust Specific Fuel Consumption 14 SSAS UAV Scientific Meeting & Exhibition, Jeddah, Saudi Arabia (June 6, 2006) Paper No. SSAS-2006-051
For the present project the propeller selected should have maximum efficiency close to a speed of 200 km/hr. Figure 12 illustrates the required TSFC variation with propulsion efficiency for airspeed of 220 km/hr and SFC of 0.33 kg/hp.hr. From Figure 12 it is clear that to meet the endurance requirements the propulsion efficiency during cruise condition should be ηp > 45 %. The recommended efficiency of propeller in the present study is 50% or better at cruise condition. This is consider a reasonable estimate compared to low speed propellers.13,14 Altitude and flight speed will change the propeller size and configuration to meet the efficiency requirement. So, more than one propeller type should be used to meet different altitudes and flight speed for optimum performance. The propulsion requirements are as the following: a) Engine :
− − − − − − − − b) Propeller: : − − − −
20 -25 hp Two stroke engine with: SFC < 0.33 kg/hp.hr (at cruise) Power to weight ratio > 2.5 kg/hp (including generator) Ability to operate to temperatures up to 45 oC Ability to operate in dusty environment (equipped with intake filter) Able to with stand axial acceleration > 10 g. Operate efficiently at max altitude 18,000 ft Self priming ignition system with min. 100 W generator. Pusher type Meets engine power 70-80 cm diameter, minimum possible size Efficiency ≥ 50% at cruise condition
III.
Conclusion
The major objective of the multi-role UAV is to be capable of long endurance (8-12 hrs) flights with a maximum payload of 10 kg. The maximum expected cruise speed is 220 km/hr, and the ceiling is 18,000 ft. This would allow the UAV to perform over 1,700 km of non-stop flight with maximum payload and without refueling. Compliance to specifications limits of different systems is expected to assure achievement to the objectives of this project. The present study was based on available information in the literature and published specifications of manufacturers. Design constrains are identified as: − Available engine performance and specifications − Available control system specifications (size – weight-power) − Communication System Other major specifications can be met through careful design and execution. Based on design sensitivity analysis the UAV is expected to have a conventional aerodynamic configuration with pusher propeller driven by a 2 stroke IC engine with > 20 hp output for improved reliability. Special heat sinks for engine heads are recommended for improved reliability of the system during hot local season. Also, the structure should be made of light 15 SSAS UAV Scientific Meeting & Exhibition, Jeddah, Saudi Arabia (June 6, 2006) Paper No. SSAS-2006-051
composite material to comply with limited aircraft dry weight of 36 kg. This would provide a tolerance of 4 kg for the structural components of the aircraft in the design and construction phase compared to the typical design in table 3. Wings should be separable for ease of store and transportation. Moreover, aerodynamic performance (L/D) have to be grater than 13, which requires a retractable landing gear configuration. In addition, the engine has to have high specific power > 2.5 hp/kg, with low SFC< 0.33 kg/hp.hr. Also, the propeller should have high efficiency during cruise (> 50%). Survivability of the UAV is improved by a recovery system composed of a parachute and an airbag in case of emergencies.
Acknowledgment The author acknowledges the support of King Fahd University of Petroleum and Minerals during this research.
References 1
Johnson, E. N., Hart, M. G., and Christophersen, H.B., “Development of an Autonomous Aerial Reconnaissance System at Georgia Tech,” AIAA's 1st Technical Conference and Workshop on Unmanned Aerospace Vehicles, Systems, Technologies, and Operations, 20-23 May 2002 2 Hennessey, C., “Autonomous Control of a Scale Airplane,” School of Engineering Science, Simon Fraser University, 14 April 2000 3 Wang K.C." UAV Design Activities in a University Environment" presented at the 9th Australian International Aerospace Congress, Canberra, Australia, 6-8 March 2001 (PDF 4.0 file) 4 Amro M. Al-Qutub and Sami El Ferik "System Definition for Unmanned Aircraft Development" KFUPM Internal Report Jun, 2003. 5 BAI Aero systems. Payloads and Sensors. www.baiaerosystems.com/payloads.html 6 UAV forum: http://www.uavforum.com/index.shtml 7 Association for Unmanned Vehicle Systems International: http://www.auvsi.org/ 8 UAV Engines Ltd.: http://www.uavenginesltd.co.uk 9 Zanzottera Engines: http://www.zazoterteraengines.com 10 Limbach Engines: http://www.limflug.de 11 3W Modellmotoren GmBH Engines: http://www.3w-modellmotoren.com 12 Hill and Peterson "Mechanics and Thermodynamics of Propulsion", Addition-Wesley, 2nd edition, 1992 13 Shaila L. Jaszlics and Rianer Stemme " Possibility of the Next Generation of High-Altitude LongEndurance Unmanned Aerial Vehicles" Technical Publication/ HALE-UAV. 14 F 2 Craig D. Paxton, Peter J. Gryn, Eerisa Hines, Ulises Perez, and Ge-Cheng Zha " High Efficiency Foreword Swept Propellers at Low Speed" AIAA 2003-1069.
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