Category AIRCRAF DESIGN

At low speeds, the maximum load factor is constrained by the aircraft maximum CL. The low-speed limit in a V-n diagram is established at the velocity at which the aircraft stalls in an acceleration flight load of n until it reaches the limit-load factor. At higher speeds, the maneuver-load factor may be restricted to the limit­load factor, as specified by the regulatory agencies.

Let VS1 be the stalling speed at 1 g. Then:

= (ovCm;) (!) or L=W=(0V S)Cim“

Let Vsn be the stalling speed at ng, where n is a number. Then:

nW = (0.5p V2„S)CLmax

Using Equations 5.1 and 5.2,

n X (0.5p VlS)CLmax = (0.5p V?„S)Clmax

or

n = Vl/Vl = until n reaches the limit-load factor (5.5)

VA is the speed at which the positive-stall and maximum-load factor limits are simul­taneously satisfied (i. e., VA = VSu/nlimit).

The negative side of the boundary can be estimated similarly.

5.7.1 High-Speed Limit

VD is equal to the maximum design speed. It is limited by the maximum dynamic pressure that an airframe can withstand. At high altitude, VD may be limited by the onset of high-speed flutter.

This is the required maneuver load factor at all speeds up to VC. (The next section defines speed limits.) Maximum elevator deflection at VA and pitch rates from VA to VD also must be considered. Table 5.1 gives the g-limit of various aircraft classes.

For military aircraft applications, in general, the factor of safety equals 1.5 but can be modified through negotiation (see Military Specifications MIL-A-8860, MIL­A-8861, and MIL-A-8870).

Typical g-levels for various types of aircraft are shown in Table 5.2. These limits are based on typical human capabilities.

5.6.1 Speed Limits

The V-n diagram (see Figure 5.2) described in Section 5.7 uses various speed limits, defined as follows:

VS: Stalling speed at normal level flight.

VA: Stalling speed at limit load. In a pitch maneuver, an aircraft stalls at a higher speed than the VS. In an accelerated maneuver of pitch­ing up, the angle of attack, a, decreases and therefore stalls at higher speeds. The tighter the maneuver, the higher is the stalling speed until it reaches VA.

VB: Stalling speed at maximum gust velocity. It is the design speed for maxi­mum gust intensity VB and is higher than VA.

VC: Maximum level speed.

VD: Maximum permissible speed (occurs in a dive; also called the placard speed).

An aircraft can fly below the stall speed if it is in a maneuver that compensates loss of lift or if the aircraft attitude is below the maximum angle of attack, amax, for stalling.

5.4 V-n Diagram

To introduce the V-n diagram, the relationship between load factor, n, and lift coef­ficient, CL, must be understood. Pitch-plane maneuvers result in the full spectrum of angles of attack at all speeds within the prescribed boundaries of limit loads. Depending on the direction of pitch-control input, at any given aircraft speed, posi­tive or negative angles of attack may result. The control input would reach either the CLmax or the maximum load factor n, whichever is the lower of the two. The higher the speed, the greater is the load factor, n. Compressibility has an effect on the V-n diagram. In principle, it may be necessary to construct several V-n diagrams repre­senting different altitudes. This chapter explains only the role of the V-n diagram in aircraft design.

Figure 5.2 represents a typical V-n diagram showing varying speeds within the specified structural load limits. The figure illustrates the variation in load factor with airspeed for maneuvers. Some points in a V-n diagram are of minor interest to con­figuration studies – for example, at the point V = 0 and n = 0 (e. g., at the top of the vertical ascent just before the tail slide can occur). The points of interest are explained in the remainder of this section.

Inadvertent situations may take aircraft from within the limit-load boundaries to conditions of ultimate-load boundaries (see Figure 5.2).

Limit load is defined as the maximum load that an aircraft can be subjected to in its life cycle. Under the limit load, any deformation recovers to its original shape and would not affect structural integrity. Structural performance is defined in terms of stiffness and strength. Stiffness is related to flexibility and deformations and has implications for aeroelasticity and flutter. Strength concerns the loads that an air­craft structure is capable of carrying and is addressed within the context of the V-n diagram.

To ensure safety, a margin (factor) of 50% increase (civil aviation) is enforced through regulations as a factor of safety to extend the limit load to the ultimate load. A flight load exceeding the limit load but within the ultimate load should not cause structural failure but could affect integrity with permanent deformation. Aircraft are equipped with g-meters to monitor the load factor – the n for each sortie – and, if exceeded, the airframe must be inspected at prescribed areas and maintained by prescribed schedules that may require replacement of structural components. For example, an aerobatic aircraft with a 6-g-limit load will have an ultimate load of 9 g. If an in-flight load exceeds 6 g (but is below 9 g), the aircraft may experience permanent deformation but should not experience structural failure. Above 9 g, the aircraft would most likely experience structural failure.

The factor of safety also covers inconsistencies in material properties and man­ufacturing deviations. However, aerodynamicists and stress engineers should cal­culate for load and component dimensions such that their errors do not erode the factor of safety. Geometric margins, for example, should be defined such that they add positively to the factor of safety.

ultimate load = factor of safety x limit load

For civil aircraft applications, the factor of safety equals 1.5 (FAR 23 and FAR 25, Vol. 3).

In steady-level flight, an aircraft is in equilibrium; that is, the lift, L, equals the air­craft weight, W, and the thrust, T, equals drag, D. During conceptual design, when generating the preliminary aircraft configuration, it is understood that the wing pro­duces all the lift with a spanwise distribution (see Section 3.14).

In equation form, for steady-level flight:

Vfei£it

Figure 5.1. Equilibrium flight

5.5.1 Load Factor, n

Newton’s law states that change from an equilibrium state requires an additional applied force; this is associated with some form of acceleration, a. When applied in the pitch plane, the force appears as an increment in lift, AL, and it would over­come the weight, W, to an increased altitude initiated by rotation of the aircraft (Figure 5.1).

From Newton’s law:

AL = centrifugal acceleration x mass = a x W/g (5.2)

The resultant force equilibrium gives:

L + AL = W + a x W/g = W(1 + a/g) (5.3)

where L is the steady-state lift equaling weight, W load factor, n, is defined as:

n = (1 + a/g) = L/ W + A L/ W = 1 + AL/ W (5.4)

The load factor, n, indicates the increase in force contributed by the centrifugal acceleration, a. The load factor, n = 2, indicates a twofold increase in weight; that is, a 90-kg person would experience a 180-kg weight. The load factor, n, is loosely termed as the g-load; in this example, it is the 2-g-load.

A high g-load damages the human body, with the human limits of the instanta­neous g-load higher than for continuous g-loads. For a fighter pilot, the limit (i. e., continuous) is taken as 9 g; for the civil aerobatic category, it is 6 g. Negative g-loads are taken as half of the positive g-loads. Fighter pilots use pressure suits to control blood flow (i. e., delay blood starvation) to the brain to prevent “blackouts.” A more inclined pilot seating position reduces the height of the carotid arteries to the brain, providing an additional margin on the g-load that causes a blackout.

Because they are associated with pitch-plane maneuvers, pitch changes are related to changes in the angle of attack, a, and the velocity, V. Hence, there is variation in CL, up to its limit of CLmax, in both the positive and negative sides of the wing incidence to airflow. The relationship is represented in a V-n diagram, as shown in Figure 5.2. Atmospheric disturbances are natural causes that appear as a gust load from any direction. Aircraft must be designed to withstand this unavoid­able situation up to a statistically determined point that would encompass almost all-weather flights except extremely stormy conditions. Based on the sudden excess in loading that can occur, margins are built in, as explained in the next section.

Aircraft speed V)

Loads on the ground are taken up by the undercarriage and then transmitted to the aircraft main structure. Landing-gear loads depend on the specification of Vstall, the maximum allowable sink speed rate at landing, and the MTOM. This is addressed in greater detail in Chapter 7, which discusses undercarriage layout for conceptual study.

5.4.1 In Flight

In-flight loading in the pitch plane is the main issue considered in this chapter. The aircraft structure must be strong enough at every point to withstand the pressure field around the aircraft, along with the inertial loads generated by flight maneuvers. The V-n diagram is the standard way to represent the most severe flight loads that occur in the pitch plane (i. e., X-Z plane), which is explained in detail in Section 5.7. The load in other planes is not discussed herein.

The aileron-induced motion generates the roll maneuver with angular velocity, p, about the X-axis, in addition to velocities in the Y-Z plane. Aircraft structures designed to the pitch-plane loading are the most critical; therefore, roll-plane load­ing is not discussed herein.

5.3.2 Yaw Plane (Z-X) Maneuver (Rudder-Induced)

The rudder-induced motion generates the yaw (coupled with the roll) maneuver with angular velocity, r, about the Z-axis, in addition to linear velocities in the Z-X plane. Aerodynamic loading of an aircraft due to yaw is also necessary for structural design.

5.3 Aircraft Loads

An aircraft is subject to load at any time. The simplest case is an aircraft stationary on the ground experiencing its own weight. Under heavy landing, an aircraft can experience severe loading, and there have been cases of structural collapse. Most of these accidents showed failure of the undercarriage, but breaking of the fuselage also has occurred. In flight, aircraft loading varies with maneuvers and/or when gusts are encountered. Early designs resulted in many structural failures in flight.

Although throttle-dependent linear acceleration would generate flight load in the direction of the flight path, pilot-induced control maneuvers could generate the extreme flight loads that may be aggravated by inadvertent atmospheric con­ditions. Aircraft weight is primarily determined by the air load generated by maneuvers in the pitch plane. Therefore, the associated V-n diagram described in Section 5.7 is useful information for proposing candidate aircraft configura­tions. Section 3.6 describes the six deg of freedom for aircraft motions – three lin­ear and three angular. Given herein are the three Cartesian coordinate planes of interest.

5.3.1 Pitch Plane (X-Z) Maneuver (Elevator/Canard-Induced)

The pitch plane is the symmetrical vertical plane (i. e., X-Z plane) in which the elevator/canard-induced motion occurs with angular velocity, q, about the У-axis, in addition to linear velocities in the X-Z plane. Changes in the pitch angle due to angular velocity q results in changes in CL. The most severe aerodynamic loading occurs in this plane.

At the initial development phase of stall (or during extreme maneuvers), airflow over the wing becomes unsteady; the separation line over the wing (or over any other lifting surface) keeps fluctuating. This causes the aircraft to shudder and is a warning to the pilot. The aircraft structure is not affected and is not necessarily at its maximum loading.

5.2.1 Flutter

This is the vibration of the structure – primarily the wing but also any other compo­nent depending on its stiffness. At transonic speed, the load on the aircraft is high while the shock-boundary layer interaction could result in an unsteady flow causing vibration over the wing, for example. The interaction between aerodynamic forces and structural stiffness is the source of flutter. A weak structure enters into flutter; in fact, if it is too weak, flutter could happen at any speed because the deformation would initate the unsteady flow. If it is in resonance, then it could be catastrophic – such failures have occurred. Flutter is an aeroelastic phenomenon.

5.1 Overview

Aircraft structures must withstand the imposed load during operations; the extent depends on what is expected from the intended mission role. The bulkiness of the aircraft depends on its structural integrity to withstand the design load level. The heavier the load, the heavier is the structure; hence, the MTOW affecting aircraft performance. Aircraft designers must comply with mandatory certification regula­tions to meet the minimum safety standards.

This book does not address load estimation in detail but rather continues with design information on load experienced by aircraft. Although the information pro­vided herein is not directly used in configuring aircraft, the knowledge and data are essential for understanding design considerations that affect aircraft mass (i. e., weight). Only the loads and associated V-n diagram in symmetrical flight are dis­cussed herein. It is assumed that designers are supplied with aircraft V-n diagrams by the aerodynamics and structures groups. Estimation of load is a specialized sub­ject covered in focused courses and textbooks. However, this chapter does outline the key elements of aircraft loads. Aircraft shaping dictates the pattern of pressure distribution over the wetted surface that directly affects load distribution. There­fore, aircraft loads must be known early enough to make a design “right the first time.”

5.1.1 What Is to Be Learned?

This chapter covers the following topics:

Introduction to aircraft load, buffet, and flutter Flight maneuvers Aircraft load

Theory and definitions (limit and ultimate load) Limits (load limit and speed limit)

V-n diagram (the safe flight envelope)

Gust envelope

5.1.2 Coursework Content

This chapter provides the basic information required to generate conceptual air­craft configurations. To continue, it is recommended that readers peruse this chapter even though there is no coursework involved yet. The chapter can be skipped if the subject has been learned in other coursework. However, readers should be able to draw schematically a representative V-n diagram of their aircraft (explained in Section 5.8).

5.2 Introduction

Loads are the external forces applied to an aircraft – whether static or dynamic, in flight or on the ground. In-flight loads are due to symmetrical flight, unsymmet­rical flight, or atmospheric gusts from any direction; on-ground loads result from ground handling and field performance (e. g., takeoff and landing). Aircraft design­ers must be aware of aircraft loads given that configurations must be capable of with­standing them. During the design study phase, aerodynamicists compute in-flight aerodynamic loads and relate the information to stress engineers, who ensure struc­tural integrity. Computation of aerodynamic load is involved, currently undertaken using computers. The subject matter concerns interaction between aerodynamics and structural dynamics (i. e., deformation occurring under load), a subject that is classified as aeroelasticity. Even the simplified assumption of an aircraft as an elastic body requires study beyond the scope of this book. Generally, conceptual design addresses rigid aircraft.

User specifications define the maneuver types and speeds that influence aircraft weight (i. e., MTOM), which then dictates aircraft-lifting and control surface design. In addition, enough margin must be allocated to cover inadvertent excessive load encountered through pilot induced maneuvers (i. e., inadvertent internal input in excess of the specifications), or sudden severe atmospheric disturbances (i. e., exter­nal input), or a combination of the two scenarios. The limits of these inadvertent sit­uations are derived from historical statistical data and pilots must avoid exceeding the margins. To ensure safety, governmental regulatory agencies have intervened with mandatory requirements for structural integrity. Load factor (not to be con­fused with the passenger load factor, as described in Section 4.4.1) is a term that expresses structural-strength requirements. The structural regulatory requirements are associated with V-n diagrams, which are explained in Section 5.7. Limits of the margins are set by the regulatory agencies. In fact, they not only stipulate the load limits, they also require mandatory strength tests to determine ultimate loads. The ultimate load tests must be completed before the first flight, with the exceptions of homebuilt and experimental categories of aircraft.

Civil aircraft designs have conservative limits; there are special considerations for the aerobatic category aircraft. Military aircraft have higher limits for hard maneuvers, and there is no guarantee that under threat, a pilot would be able to adhere to the regulations. Survivability requires widening the design limits and strict maintenance routines to ensure structural integrity. Typical human limits are currently taken at 9 g in sustained maneuvers and can reach 12 g for instanta­neous loading. Continuous monitoring of the statistical database retrieved from aircraft-mounted “black boxes” provides feedback to the next generation of aircraft design or at midlife modifications. A g-meter in the flight deck records the g-force and a second needle remains at the maximum g reached in the sortie. If the pre­scribed limit is exceeded, then the aircraft must be grounded for a major inspection and repaired, if required.

An important aspect of design is to know what could happen at the extreme points of the flight envelope (i. e., the V-n diagram). In the following sections, buffet and flutter are introduced.

This extended section of the book can be found on the Web at www. cambridge .org/Kundu and gives introductory comments on typical military aircraft classifi­cation; military aircraft role, statistics, and design considerations; and some rela­tively newer requirements (evolutionary patterns), and so forth. Figure 4.30 shows (a) Lockheed F104, Starfighter; (b) McDonnell F4, Phantom; (c) Grumman F14, Tomcat; (d) Northrop F117; and (e) Lockheed F22.

Figure 4.30. Chronology of fighter aircraft design evolution (USA)

4.9 Military Aircraft Mission

This extended section of the book can be found on the Web at www. cambridge .org/Kundu and describes military aircraft multiroles, indicating that the same class

of military aircraft can have a wide variety of payload ranges. Figure 4.31 shows weapon configurations for (a) air interdiction, (b) close air support, (c) air defense, and (d) maritime attack.

Figure 4.31. Typical multirole missions

4.10 Military Aircraft Statistics (Sizing Parameters – Regression Analysis)

This extended section of the book can be found on the Web at www. cambridge .org/Kundu and gives the statistics of military aircraft as discussed in the following subsections.

4.10.1 Military Aircraft Maximum Take-off Mass (MTOM) versus Payload

In this subsection, at www. cambridge. org/Kundu, Figure 4.32 shows typical statistics of military aircraft payload – range.

Figure 4.32. Military aircraft payload – range (no drop tank or refueling)

4.10.2 Military MTOM versus OEM

In this subsection, at www. cambridge. org/Kundu, Figure 4.33 gives the relation between MTOM and OEM, as well as the operational empty mass fraction (ratio of OEM to MTOM).

Figure 4.33. MTOM versus OEM

4.10.3 Military MTOM versus Fuel Load Mf

In this subsection, at www. cambridge. org/Kundu, Figure 4.34 gives the relationship between internal fuel load and fuel fraction versus MTOM.

Figure 4.34. MTOM versus fuel load

4.10.4 MTOM versus Wing Area (Military)

In this subsection, at www. cambridge. org/Kundu, Figure 3.35 shows wing area, SW, and wing-loading MTOM/SW versus MTOM.

Figure 4.35. MTOM versus wing area

4.10.5 MTOM versus Engine Thrust (Military)

In this subsection, at www. cambridge. org/Kundu, Figure 4.36 presents the rela­tionship between total Tsls and the two types of aircraft mass (e. g., MTOM and TTOM).

Figure 4.36. Aircraft weight versus total take-off thrust

4.10.6 Empennage Area versus Wing Area (Military)

This subsection, at www. cambridge. org/Kundu, gives a brief comment on military aircraft empennage area.

4.10.7 Aircraft Wetted Area versus Wing Area (Military)

This brief subsection, at www. cambridge. org/Kundu, is on military aircraft wetted and wing areas.

4.11 Military Aircraft Component Geometries

This extended section of the book can be found on the Web at www. cambridge .org/Kundu and describes military aircraft component geometries (e. g., fuselage group, wing group, empennage group, and Nacelle group/intake).

4.12 Fuselage Group (Military)

This extended section of the book can be found on the Web at www. cambridge .org/Kundu and describes military aircraft fuselage group.

4.13 Wing Group (Military)

This extended section of the book can be found on the Web at www. cambridge .org/Kundu and describes the military aircraft wing group with Figure 4.37. Military trainer-aircraft wing group is illustrated in Figure 4.38.

Figure 4.37. Fighter aircraft configurations

One surface configuration: (Figure 4.37a – Mirage 2000 and SAAB Draken).

Two surface configuration: (Figure 4.37b – MIG 21 and Mirage F1).

(Figure 4.37c – Eurofighter and SAAB Viggen) (Figure 4.37d – F16 and F18)

Three surface configuration: (Figure 4.37e – SU 37 and SU 47).

Figure 4.38. Advanced jet trainer aircraft capable of close support combat

4.13.1 Generic Wing Planform Shapes

This extended subsection, at www. cambridge. org/Kundu, describes how military air­craft wing planforms can be presented in a unified manner and includes civil designs (i. e., from delta to rectangular shapes) as shown in Figure 4.39.

Figure 4.39. Wing planform shape

4.14 Empennage Group (Military)

This extended section of the book can be found on the Web at www. cambridge .org/Kundu and describes various aspects of military aircraft empennage configurations and available options using several figures (listed above) and Figure 4.40 (YF12, F29 and B2).

Figure 4.40. Empennage options

4.15 Intake/Nacelle Group (Military)

This extended section of the book can be found on the Web at www. cambridge .org/Kundu and gives a broad classification of military fighter aircraft engine-intake configuration, as given in Chart 4.3. Some older design engine positions are shown in Figure 4.41 (P38, B & V141, Heinkel 162, F107, Corsair, and a Tupolev design).

Chart 4.3. Types of empennage configurations

Figure 4.41. Options for engine positions of some older designs

4.16 Undercarriage Group

This extended section of the book can be found on the Web at www. cambridge .org/Kundu and refers to Chapter 7.

4.17 Miscellaneous Comments

This extended section of the book can be found on the Web at www. cambridge .org/Kundu and gives some pertinent comments on miscellaneous aspects of military aircraft design.

4.18 Summary of Military Aircraft Design Choices

This extended section of the book can be found on the Web at www. cambridge .org/Kundu and summarizes military aircraft design choices and various approaches to it.