CONCEPTUAL DESIGN OF FLYING VEHICLE

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1 International Journal of Mechanical Engineering and Technology (IJMET) Volume 8, Issue 6, June 2017, pp , Article ID: IJMET_08_06_049 Available online at ISSN Print: and ISSN Online: IAEME Publication Scopus Indexed CONCEPTUAL DESIGN OF FLYING VEHICLE Alka Sawale Assistant Professor, Department of Aeronautical Engineering, MLR Institute of Technology, Hyderabad, India Sreekanth Sura Assistant Professor, Department of Aeronautical Engineering, MLR Institute of Technology, Hyderabad, India Anitha D Assistant Professor, Department of Aeronautical Engineering, Institute of Aeronautical Engineering, Hyderabad, India B. Subbaratnam Professor and Head, Department of Mechanical Engineering, Vardhaman College of Engineering, Hyderabad, India ABSTRACT The primary objective of this report is to design a solar powered unmanned aerial vehicle with less weight of around 3kg.Another objective of this report is to provide an initial selection of the solar powered UAV. A comparative study of other solar powered will be done to have configurations of this solar powered UAV. For initial configuration there will be a weight estimation for this UAV, initial selection of weight distribution will be discussed. Since Power is most crucial parameter for solar powered UAV, so will be looking at fundamental equations of power. Finally a drag polar estimation will be done. Key words: Unmanned Aerial Vehicle (UAV), solar power, weight estimation, weight distribution. Cite this Article: Alka Sawale, Sreekanth Sura, Anitha D and B. Subbaratnam. Conceptual Design of Flying Vehicle. International Journal of Mechanical Engineering and Technology, 8(6), 2017, pp INTRODUCTION A flying wing is sometimes represented as theoretically the most aerodynamically efficient (lowest drag) design configuration for a fixed wing aircraft. It also would offer high structural efficiency for a given wing depth, leading to light weight and high fuel efficiency. Because it lacks conventional stabilizing surfaces and the associated control surfaces, in its purest form editor@iaeme.com

2 Conceptual Design of Flying Vehicle the flying wing suffers from the inherent disadvantages of being unstable and difficult to control. These compromises are difficult to reconcile, and efforts to do so can reduce or even negate the expected advantages of the flying wing design, such as reductions in weight and drag. Moreover, solutions may produce a final design that is still too unsafe for certain uses, such as commercial aviation. Further difficulties arise from the problem of fitting the pilot, engines, flight equipment, and payload all within the depth of the wing section. Other known problems with the flying wing design relate to pitch and yaw. Pitch issues are discussed in the article on tailless aircraft. In some flying wing designs, any stabilizing fins and associated control rudders would be too far forward to have much effect, thus alternative means for yaw control are sometimes provided. One solution to the control problem is differential drag: the drag near one wing tip is artificially increased, causing the aircraft to yaw in the direction of that wing. Typical methods include. A consequence of the differential drag method is that if the aircraft maneuvers frequently then it will frequently create drag. So flying wings are at their best when cruising in still air: in turbulent air or when changing course, the aircraft may be less efficient than a conventional design. Figure 1 Bi-directional flying wing. Top-down view The supersonic bi-directional flying wing design comprises a long-span low speed wing and a short-span high speed wing joined in the form of an unequal cross. The proposed craft would take off and land with the low-speed wing across the airflow, then rotate a quarter-turn so that the high-speed wing faces the airflow for supersonic travel has funded a study of the proposal The design is claimed to feature low wave drag, high subsonic efficiency and little or no sonic boom. The proposed low-speed wing would have a thick, rounded airfoil able to contain the payload and a long span for high efficiency, while the high-speed wing would have a thin, sharp-edged airfoil and a shorter span for low drag at supersonic speed. The Defense Advance Research Academy (DARA) has researched a solar powered HALE UAV. The main idea behind the project called vulture is to combine the key benefits of both an aircraft and a satellite into one system and to keep these systems in air continuously for 5 years. If the analysis is successful we will undergo the process of making prototype of this UAV with some advanced features included in it editor@iaeme.com

3 Alka Sawale, Sreekanth Sura, Anitha D and B. Subbaratnam 2. MISSION SPECIFICATION 2.1. Mission Profile The mission profile for this mission is shown as:- Figure 2 Mission Profile Table 1 Dimensions WING SPAN 280 ROOT CHORD 80 TIP CHORD 50 SWEEP ANGLE degrees WING AREA Swept Wings Neutral Point and Stability We have already learned that the center of gravity must be located in front of the neutral point. While the n.p. of an un swept, rectangular wing is approximately at thec/4point, the n.p. of a swept, tapered wing must be calculated. The following procedure can be used for a simple, tapered and, swept wing. First, we Calculate the mean aerodynamic chord length from the sweep angle: of a tapered wing, which is independent With the root chord lr, the tip chord lt and the tap. We can also calculate the span wise location of the mean Chord, using the span b, The n.p. of our swept wing can be found by drawing a line, parallel to the fuselage center line, at the spanwise station y. The chord at this station should be equal to. The n.p. is approximately located at the c/4 point of this chord line (see the sketch below) editor@iaeme.com

4 Conceptual Design of Flying Vehicle Figure 3 Geometric parameters of a tapered, swept wing Instead of using the graphical approach, the location of the neutral point can also be calculated by using one of the following formulas, depending on the taper ratio:, if taper ratio> 0.375, if taper ratio< The c.g must be placed in front of this point, and the wing may need some twist (washout) to get a sufficiently stable wing Finding the Required Twist ßreq Using graph 1, we enter the graph with the aspect ratio AR on the horizontal axis, and draw a vertical line upwards, until we intersect the curve, corresponding to the sweep angle of the c/4 line. Continuing to the axis on the left border, we find the standard value β*req for the required twist angle. This standard value is valid for a wing, which is trimmed at = 1.0 and has a stability coefficient of β* =10% (see above), and Uses airfoils with a moment coefficient of zero. From the standard value we calculate the true, required twist angle, using the formula inset into the graph. Therefore, we calculate the ratio of our target lift coefficient to the standard lift coefficient (CL/ ) and the ratio of our desired stability coefficient to the standard. We see, that a reduction of the lift coefficient to CL=0.5 also reduces the required twist by 50%. Also, if we use a smaller stability margin β, we need a smaller amount of twist. Figure 4 Finding the Required Twist Graphs editor@iaeme.com

5 Alka Sawale, Sreekanth Sura, Anitha D and B. Subbaratnam Variation of zero lift angle If we use different airfoils at root and tip, they may have different zero lift directions, which influences the equilibrium state. The geometric twist has to be reduced by the difference of the zero lift directions β 0 of tip and root sections: Using the same airfoil for both sections, we can set β0 to zero Influence of the Airfoil Moment coefficients The moment coefficient of the airfoils contributes to the equilibrium, and has to be taken into account for the calculation of the twist. Graph 2 can be used to find the equivalent twist due to the contribution of Cm, which has to be subtracted from the required twist. If we use airfoils with positive moment coefficients, the contribution will be positive, which results in a reduction of the amount twist, highly cambered airfoils yield negative values βcm, which force us to build more twist into the wing. Similar to the previous graph, We enter with the aspect ratio, intersect with the sweep curve and read the value for βcm from the left hand axis. Figure 5 Finding the additional twist due to the airfoils moment coefficient. Again, the graph has been plotted for a certain standard condition, which is a moment coefficient of cm* = 0.05 (note: positive value). We apply the ratio of the moment coefficients (cm/cm*) to find the contribution βcm of the moment coefficient to the geometric twist. This contribution has to be subtracted from the required twist angle, too. Using the usual, cambered airfoils with negative moment coefficients will change the sign of the ratio cm/cm*, which results in negative β Cm values. This means, that the subtraction from βreq will actually be an addition, increasing the geometric twist angle. If we have different airfoils at root and tip, we can use the mean moment coefficient (cm,tip+ cm,root)/2 to calculate the ratiocm/cm*.finally, we can calculate the geometric twist angle βgeo, which has to be built into the wing: S = (l_r + l_t)/2 * b = m² And the aspect ratio AR = b²/s = 11.0 And the mean moment coefficient cm = (cm,r + cm,t)/2 = editor@iaeme.com

6 Conceptual Design of Flying Vehicle Using graph 1, we find β*req = 11.8, which has to be corrected to match our design lift coefficient and the desired stability margin: 11.8 * (0.5/1.0) * (0.05/0.1) = 2.95 This means that our model would need a twist angle of 2.95 (wash out) from rot to tip, if we would use a symmetrical airfoil section. The difference of the zero lift angle of tip and root section is Now we read the twist contribution of the moment coefficient from graph 2, which is β*cm = 5.8, which has to be corrected for our smaller mean moment coefficient: 5.8 * (0.02/0.05) = 2.32 Finally, we calculate the geometric twist from = The negative value means, that we could use a small amount of wash-in! This is because we have already enough stability due to the selection of airfoils with reflexes camber lines. Since the calculated amount is very small, we can use the same angle of incidence for the root and tip ribs. Since the presented method is not perfect, we can assume accuracy to 1 degree, which is also a reasonable assumption for the average building skills. 3. OVERALL WING AND AIRFOIL CONFIGURATION Based on Roskam, conventional configuration is used for the HALE-SUPAV. The UAV will not have higher range and thus will preliminary fly on land and if necessary on water. Adding another alternative fuel system would give more rang but will also increase complexity and weight, which is not recommended. The geometry of the wing should have negligible sweep because the aircraft will be operating at low speeds. Sweep will also increase weight and reduce available solar cell area, both of which will hinder the aircraft s performance. An initial airfoil selection will be the Selig 1223, as shown in Figure. This airfoil has12.14% maximum thickness-to-chord ratio at roughly 20% from the leading edge. The Selig 1223 airfoil was chosen as the initial configuration because it has all the characteristics, which requires in solar power high altitude and long endurance airplanes. The first important one is that it a low Reynolds number will be generated throughout the mission, and therefore an airfoil that has ideal characteristics at low speed has been chosen. High lift to drag ratio is also one of the important characteristics. Using XFLR5 software that analyses the airfoils, a graph of L/D vs. angle of attack was created and is shown in Figure 8. A legend that is used for the different Reynolds number used is shown in figure 5. As the figures show, not only does this airfoil have a high lift-to-drag ratio, but also it has a fairly wide operating angle of attack where the lift-to-drag ratio is optimum. Other airfoil characteristics are shown in figures editor@iaeme.com

7 Alka Sawale, Sreekanth Sura, Anitha D and B. Subbaratnam Figure 6 Selig 1223 Airfoil 4. ANALYSIS Figure 7 Plot Points Figure 8 CL/CD Vs ALPHA Figure 9 CL-ALPHA Curve Figure 10 CM-ALPHA Curve Figure 11 CL/CD Curve editor@iaeme.com

8 Conceptual Design of Flying Vehicle The drag polar will now be calculated for the airplane itself. Using the value of the calculated zero-lift drag coefficient, the overall aircraft drag coefficient can be calculated using. Assuming a lift coefficient of from Section 8.3, as well as an Oswald efficient factor of 0.9, the total aircraft drag coefficient can be found as This will be used to find lift to drag ratio as shown: XFLR5 program also used to find out lift to drag ratio for this airplane and the data are shown in appendix F. During the analysis, three different values of drag coefficient have come out hence the lift to drag ratios. 5. RESULT AND DISCUSSION The drag polar explains in table 9 shows that lift to drag ratio ranges from 38to 52. The major difference is in analysis because other two have less difference. The analysis does not include fuselage drag coefficient as other so that analysis should be most inaccurate among others. The average value between these two approaches can be used in further analysis if needed Flying wing CG calculator Figure 12 CG Calculations Figure 13 Solid works Model Figure 14 Side View with Winglet 6. CONCLUSIONS The current desire for a greener society, an alternative source of energy for aircraft is needed. There are many alternative energy solutions that are promising including bio-fuel and hydrogen fuel cells, but nothing is as limitless as solar technology. As, mentioned throughout the project, the application of high altitude long endurance UAVs can potentially be very large, whether it is in weather surveillance, studying natural disaster, or fire direction. The solar power UAV design discussed weight, has a large wingspan of 280mm, and hold upto 300grams of payload, which is more than enough for all the surveillance and autopilot instruments. The advances in solar technology have made it so the concept of solar powered UAVs and MAVs is not just a theory anymore. Solar power airplanes are necessary for greener society and can be an important part of the future of aviation. Hence we consider this UAV for future scope of making it a solar powered UAV editor@iaeme.com

9 Alka Sawale, Sreekanth Sura, Anitha D and B. Subbaratnam REFERENCES [1] Bilstein, RE. Air travel and the traveling public: the American experience, In From airships to airbus: the history of civil and commercial aviation. Vol. 2, Pioneers and operations (ed. W.F. Trimble).Proc. International Conference on the History of Civil and Commercial Aviation, Lucerna (CH), 1992, Smithsonian Institution, Washington (USA), [2] Liebeck, RH, Page, MA and Rawdon, BK. Blended Wing-Body Subsonic Commercial Transport. 36thAerospace Science Meeting & Exhibit, Reno (NV, USA), AIAA Paper , [3] McMasters, JH and Kroo, IM. Advanced configurations for very large transport airplanes. Aircraft Design, Vol. 1, No. 4, pp , 1998 [4] Lange, RH. Review of Unconventional Aircraft Design Concepts. Journal of Aircraft, Vol. 25, No. 5, pp , [5] Global Market Forecast Airbus Industrie, Toulouse (F), [6] Martínez-Val, R, Pérez, E, Muñoz, T and Cuerno, C. Design Constraints in the Payload-Range Diagram of Ultrahigh Capacity Transport Airplanes. Journal of Aircraft, Vol. 31, No. 6, pp , [7] Nimmy George, Grid Connected PV System Using 9-Level Flying Capacitor Multilevel Inverter. International Journal Of Electrical Engineering & Technology (IJEET). 5(12), 2014, pp [8] Prithvish Mamtora, Sahil Shah, Vaibhav Shah, Vatsal Vasani, Unmanned Aerial Vehicle (UAV). International Journal of Electronics and Communication Engineering & Technology (IJECET), 4(5), 2013, pp [9] T. Prabhu, Unmanned Surface Vehicle (USV) for Coastal Surveillance. International Journal of Mechanical Engineering and Technology, 7(3), 2016, pp editor@iaeme.com