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CHAPTER 2

Design Disciplines

2.1 INTRODUCTION

There are several design disciplines which work in parallel within a UAV design project. Some examples are: (1) aerodynamic design, (2) structural design, (3) propulsion system design, (4) power transmission system design, (5) mechanical system design, and (6) control surfaces design, (7) ground station, and (8) launch and recovery system. This chapter briefly covers the first six topics disciplines; the other two are presented in Chapters 8 and 9, respectively. Due to the limited volume of the book, only the basic fundamentals are presented. The interested reader should refer to Sadraey [37] for more details. Table 2.1 shows the UAV major components and their primary functions.

Table 2.1: UAV vehicle major components and their functions

In a UAV design process, some UAV parameters must be minimized (e.g., weight), while some other variables must be maximized within constraints (e.g., range, endurance, maximum speed, and ceiling), and also others must be evaluated to ensure that they are acceptable. The optimization process must be accomplished through a systems engineering approach. In some cases, the design of the UAV may impose slight to considerable changes to the UAV mission during the conceptual design process. The strong relationship between the analysis and the influencing parameters allow definite, traceable relationships to be constructed. In the case of a UAV design, the major parameters are derived almost completely from operational and performance requirements.

It is clear that some steps may be moved along with regard to the UAV mission, design team members, past design experiences, design facility, and manufacturing technologies. As it is observed, the design process is truly an iterative process and there are several modification steps to satisfy all design requirements. An important feature of the design process is the lessons learned in the past. The lesson will be utilized in improving the next generation, for instance, the major setback during Phase II flight testing of Global Hawk (Figure 1.1) was the destruction of air vehicle 2 on March 29, 1999, during the program’s 18th sortie. The loss of air vehicle 2 and its payload was estimated at $45 million. Of more importance, however, was the fact that the program lost its only integrated sensor suite. The crash was due to a lack of proper frequency coordination between the Nellis Air Force Base and EAFB flight test ranges. Essentially, Nellis officials who were testing systems in preparation for Global Hawk’s first planned D&E exercise were unaware that Global Hawk was flying over China Lake Naval Air Weapons Station, which is within EAFB’s area of responsibility. Thus, many changes have been applied in the design of Northrop Grumman RQ-4B Global Hawk as compared with RQ-4A. For instance, the Northrop Grumman RQ-4B Global Hawk has a 50% payload increase, larger wingspan (130.9 ft) and longer fuselage (47.6 ft), and new generator to provide 150% more electrical output.

The integration of system engineering principles with the analysis-driven UAV design process demonstrates that a higher level of integrated vehicle can be attained; identifying the requirements/functional/physical interfaces and the complimentary technical interactions which are promoted by this design process. The details of conceptual design phase, preliminary design phase, and detail design phase were introduced in Chapter 1. In this chapter, design activities in several underlying disciplines are provided.

2.2 AERODYNAMIC DESIGN

The primary aerodynamic function of the UAV components (e.g., wing) is to generate sufficient lift force or simply lift (L). However, they have two other productions, namely drag force or drag (D) and nose-down pitching moment (M). While a UAV designer is looking to maximize the lift, the other two (drag and pitching moment) must be minimized. In fact, a wing is considered as a lifting surface that lift is produced due to the pressure difference between lower and upper surfaces. Aerodynamics textbooks are a good source to consult for information about mathematical techniques for calculating the pressure distribution over the wing and for determining the flow variables.

During the aerodynamic design process, several parameters must be determined. For instance, for a wing, they are as follows: (1) wing reference (or planform) area, (2) number of the wings, (3) vertical position relative to the fuselage (high, mid, or low wing), (4) horizontal position relative to the fuselage, (5) cross-section (or airfoil), (6) aspect ratio (AR), (7) taper ratio (λ), (8) tip chord (Ct), (9) root chord (Cr), (10) mean Aerodynamic Chord (MAC or C), (11) span (b), (12) twist angle (or washout) (α_t), (13) sweep angle (Λ), (14) dihedral angle (Γ), (15) incidence (i_w) (or setting angle, α_set), (16) high-lift devices such as flap, (17) aileron, and (18) other wing accessories.

One of the necessary tools in the wing design process is an aerodynamic technique to calculate wing lift, wing drag, and wing pitching moment. With the progress of the science of aerodynamics, there are variety of techniques and tools to accomplish this time consuming job. A variety of tools and software based on aerodynamics and numerical methods have been developed in the past decades. The CFD¹ Software based on the solution of Navier-Stokes equations, vortex lattice method, thin airfoil theory, and circulation are available in the market. The application of such software packages, which is expensive and time-consuming, at this early stage of wing design seems unnecessary.

Wing is a three-dimensional component, while the airfoil is a two-dimensional section. Because of the airfoil section, two other outputs of the airfoil, and consequently the wing, are drag and pitching moment. The wing may have a constant or a non-constant cross-section across the wing. There are two ways to determine the wing airfoil section, the airfoil design and the airfoil selection. The design of the airfoil is a complex and time consuming process and needs expertise in fundamentals of aerodynamics at graduate level. Since the airfoil needs to be verified by testing it in a wind tunnel, it is expensive too.

Selecting an airfoil is a part of the overall wing design. Selection of an airfoil for a wing begins with the clear statement of the flight requirements. For instance, a subsonic flight design requirements are very much different from a supersonic flight design objectives. On the other hand, flight in the transonic region requires a special airfoil that meets Mach divergence requirements. The designer must also consider other requirements such as airworthiness, structural, manufacturability, and cost requirements. In general, the following are the criteria to select an airfoil for a wing with a collection of design requirements:

1. the airfoil with the highest maximum lift coefficient ;

2. the airfoil with the proper ideal or design lift coefficient ;

3. the airfoil with the lowest minimum drag coefficient ;

4. the airfoil with the highest lift-to-drag ratio ;

5. the airfoil with the highest lift curve slope ;

6. the airfoil with the lowest (closest to zero; negative or positive) pitching moment coefficient (C_m);

7. the proper stall quality in the stall region (the variation must be gentle, not sharp);

8. the airfoil must be structurally reinforceable. The airfoil should not that much thin that spars cannot be placed inside;

9. the airfoil must be such that the cross section is manufacturable;

10. the cost requirements must be considered; and

11. other design requirements must be considered. For instance, if the fuel tank has been designated to be places inside the wing inboard section, the airfoil must allow the sufficient space for this purpose.

12. If more than one airfoil is considered for a wing, the integration of two airfoils in one wing must be observed.

In designing the high lift device for a wing, the following parameters must be determined: (1) high lift device location along the span; (2) the type of high lift device; (3) high lift device chord (C_f); (4) high lift device span (b_f); and (5) high lift device maximum deflection (down) (δ_fmax). For fundamentals of aerodynamics, please refer to references such as Anderson [44] and Shevell [45].

2.3 STRUCTURAL DESIGN

The structure of a conventional fixed-wing UAV consists of five principal units: fuselage, wings, horizontal tail, vertical tail, and control surfaces. The landing gear is also part of structure, but will be covered in Section 2.5. Engine pylon, engine inlet (for supersonic UAVs), fairings (and fillets), and landing gear bay doors are also assumed as part of aircraft structure. The primary functions of the structure is (1) to keep the aerodynamic shape of the UAV and (2) to carry the loads. Airframe structural components are constructed from a wide variety of materials. The earliest aircraft were constructed primarily of wood. Steel tubing and the most common material, aluminum, followed. Many newly certified aircraft are built from molded composite materials, such as glass/epoxy and carbon fiber.

Structural members of a fuselage mainly include stringers, longerons, bulkheads, and skin. The structural members in a wing/tail are spar, rib, stiffeners, and skin. The fuselage/wing/tails skin can be made from a variety of materials, ranging from impregnated fabric to plywood, aluminum, or composites. Under the skin and attached to the structural components are the many components that support airframe function. The entire airframe and its components are joined by rivets, bolts, screws, and other fasteners. Welding, adhesives, and special bonding techniques are also employed.

The most common form of UAV structure is semi-monologue (single shell) which implies that the skin is stressed/reinforced. The structural members are designed to carry the flight loads or to handle stress without failure. In designing the structure, every square inch of wing and fuselage, must be considered in relation to the physical characteristics of the material of which it is made. Every part of the structure be planned to carry the load which is applied on it.

The structural designer will determine flight loads, calculate stresses, and design structural elements such as to allow the UAV components to perform their aerodynamic functions efficiently. This goal will be considered simultaneously with the objective of the lowest structural weight. The most common tool in structural analysis is the use of finite element methods (FEM). One of the earliest and the most well-known computer software is NASTRAN, developed by NASA in the mid-1960s. The stress analysis is the basic calculation to determine the safety factor. There are five major stresses to which structural members are subjected: (1) tension, (2) compression, (3) torsion, (4) shear, and (5) bending. A single member of the structure is often subjected to a combination of stresses.

Fuselage usually consists of frame assemblies, bulkheads, and formers. The skin is reinforced by longitudinal members called longeron. Often, wings/tails are of full cantilever design. In general, wing construction is based on one of three fundamental designs: (1) monospar, (2) multispar, and (3) box beam. Spars are the principal structural members of the wing. They correspond to the longeron of the fuselage. Spars run parallel to the lateral axis of the aircraft, from the fuselage toward the tip of the wing, and are usually attached to the fuselage by a beam, or a truss. Generally, a wing has two spars. One spar is usually located at the maximum thickness, and the other about two-thirds of the distance toward the wing’s trailing edge (in front of flap/control-surface).

Honeycomb structured wing panels are often used in composite wings. Nacelles (i.e., pods) are streamlined enclosures used primarily to house the engine and its components. Engine mounts are also found in the nacelle. These are the assemblies to which the engine is fastened. They are usually constructed from chrome/molybdenum steel tubing in light UAV and forged chrome/nickel/molybdenum assemblies in larger UAVs. Cowling are the detachable panels covering those areas into which access must be gained regularly, such as the engine and its accessories. In the design of airframe, several factors such as ultimate load, aerodynamic loads (pressure distribution), weight loads (e.g., fuel and engine), weight distribution, gust load, load factor, propulsion loads, landing loads (e.g., brake), and aero-elasticity effects must be considered.

One of the design requirements for some military UAVs is stealth. In the concept of stealth, the three basic methods of minimizing the reflection of pulses back to a receptor are: (1) to manufacture appropriate areas of the UAV from radar-translucent material such as Kevlar or glass composite as used in radomes which house radar scanners; (2) to cover the external surfaces of the aircraft with RAM (radar absorptive material); and (3) to shape the aircraft externally to reflect radar pulses in a direction away from the transmitter. The acoustic (i.e., noise) wavelength (signature) range for detecting an air vehicle is 16 m–2 cm.

The operating flight loads limits on a UAV are usually presented in the form of a V-n diagram. Structural designers will construct this diagram with the cooperation of the flight dynamics group. The diagram will determine the structural failure areas, and area of structural damage/failure. The UAV should not be flown out of the flight envelope, since it is not safe for the structures. The UAV structural design is out of scope of this book, you may refer to references such as Megson [47] for more details.

2.4 PROPULSION SYSTEM DESIGN

A heavier-than-air craft (UAV) requires a propulsion system in order to have a sustained flight. Without a proper aero-engine or powerplant, a heavier-than-air vehicle can only glide for a short time. The contribution of a powerplant to an aircraft is to generate the most influential force in the aircraft performance; that is, the propulsive force or thrust. The secondary function of the propulsion system is to provide power/energy to other subsystems such as hydraulic system, electric system, pressure system, air conditioning system, and avionics. These subsystems rely on the engine power to operate.

Soon after the design requirements and constraints are identified and prioritized, the propulsion system designer will begin to select the type of engine. There are a number of engine types available in the market for flight operations. They include: electric (battery), solar-powered, piston-prop, turbojet, turbofan, turboprop, turboshaft, ramjet, and rocket engines.