Category AIRCRAF DESIGN

This section derives the relevant performance equations used in this book. For more details, readers may consult references [2] through [6].

Using the engine and aircraft data developed thus far during the conceptual design phase, the next section verifies whether the configured aircraft satisfies the airworthiness (i. e., FAR) and customer requirements in the takeoff/landing, the ini­tial climb rates, and the maximum initial cruise speed, as well as the payload-range capability (i. e., civil aircraft). Certifying agencies have mandatory requirements to ensure safety at takeoff and landing. Airworthiness regulations differ among coun­tries. For further details, readers may refer to the respective regulation – most of which appear in the official Web sites.

13.4.1 Takeoff

During a takeoff ground run, an aircraft under maximum thrust accelerates, gaining speed until a suitable safe speed is reached. The pilot then initiates rotation of the aircraft by gently pulling back the control stick or wheel (i. e., the elevator is going up) for a liftoff.

Designers must know the sequence of the takeoff speed schedules stipulated by the certifying agencies. To ensure safety, the agencies demand mandatory require­ments for taking off with one engine inoperative to clear a 35-ft height that repre­sents an obstacle. A one-engine inoperative TOFL is computed by considering the balanced field length (BFL) when the stopping distance after an engine failure at

the decision speed, V1, is the same as the distance taken to clear the obstacle at the MTOM (Figure 13.8). Figure 13.8 gives the various speed schedules during a takeoff run, which are explained as follows:

Vi: This is the decision speed. An engine failure below this speed would

result in an aircraft not being able to satisfy takeoff within the specified field length but able to stop. If an engine fails above the V1 speed, the aircraft should continue the takeoff operation.

Vmc: This is the minimum control speed at which the rudder is effective to con­trol the asymmetry created by a one-engine failure. It should be lower than Vi; otherwise, at the loss of one engine at Vi, an aircraft cannot be controlled if it continues the takeoff operation.

VR: This is the speed at which a pilot initiates the action to rotate an aircraft

for liftoff; it should be >1.05 Vmc. Once this is accomplished, reaching VLo and V2 occurs as an outcome of the action. VR should be more than

Vstall.

Vmu: There is a minimum “unstick” speed, above which an aircraft can be

made to lift off. The speed should be slightly above VR. In fact, Vmu determines VR. If a pilot makes an early rotation, then Vmu may not be sufficient for liftoff and the aircraft tail drags until it gains sufficient speed for liftoff.

VLO: This is the speed at which the aircraft lifts off the ground; it is closely associated with VR. If one engine is inoperative, it should have a VLO > 1.05Vmu.

V2: This is the takeoff climb speed at a 35-ft height, also known as the first-

segment climb speed; it is also closely associated with VR. FAR require that V2 = 1.2Vstall (at a minimum; it can be higher).

Vr: This is the brake-application speed with a one-engine failure (VB > V1).

The first-segment speed schedules are interrelated and expressed in terms of the ratios of Vstall, as given in Table 13.4. The velocity ratios in Table 13.4 comprise a typical range and can deviate a little as long as there is compliance with the FAR stipulation (marked with an asterisk in the table). Table 13.5 provides details of the climb segments.

Some engines at the takeoff rating have an APR that could generate a 5% higher thrust than the maximum takeoff thrust for a short period. These types of engines are not considered in this book.

The higher the thrust loading (T/W), the higher is the aircraft acceleration. For smaller changes, VR/Vstall and Vbo/Vstall may be linearly decreased with an increase in T/W. The decision speed V1 is established through iterations, as described in Section 13.5.1. In a family of derivative aircraft, the smaller variant can have a V1 close to the VR.

Table 13.5 lists the aircraft configurations and power settings for the climb seg­ments. The first – and second-segment climb schedule has FAR requirements; how­ever, the initial enroute climb capability is a customer requirement, not a FAR requirement.

Military aircraft requirements (i. e., MIL-C5011A) are slightly different than civil aircraft requirements; the first-segment clearing height is 50 ft rather than 35 ft. Many military aircraft have a single engine in which the concept of BFL is not applicable. Military aircraft must satisfy the critical field length (CFL) as described in Section 13.6.1. The second-segment rate of climb must meet a minimum of 500 ft/minute for a multiengine aircraft.

Figure 10.43a shows the uninstalled maximum climb SHP in nondimensional form and fuel consumption (sfc) up to a 30,000-ft altitude for four TAS from 50 to 200 kts. Following the same procedure as described in Section 10.11.4, the available installed thrust and fuel-flow rates at the maximum climb rating are worked out using a 4% loss of thrust due to installation effects. A sample computational table at a 30,000-ft altitude is in Table 13.2.

Figure 13.6 plots the available installed thrust and fuel flow at the maximum climb rating from sea level to a 30,000-ft altitude. The specified requirement of the initial climb rate for the example in this book is 4,500 ft/min.

Maximum Cruise Rating (Turboprop): Standard Day

Figure 10.44 shows the maximum cruise SHP in nondimensional form and fuel con­sumption (sfc) for speeds from the TAS of 50 to 300 kts and an altitude up to

Altitude ft

Figure 13.7. Thrust and fuel flow at maximum cruise rating 30,000 ft. Progressing as in the climb performance, the maximum cruise thrust and fuel-flow rate can be computed. A sample computation at a 30,000-ft altitude is in Table 13.3 using a 4% loss of thrust due to installation effects.

Figure 13.7 plots the available installed thrust and fuel flow at the maximum cruise rating. The initial maximum cruise speed in the example is 320 mph (270 kts) at a 25,000-ft altitude.

The installed takeoff thrust of a turboprop is plotted in Figure 10.40. It is repeated in this section as Figure 13.5 to keep all available thrust graphs in one section. At the takeoff rating, engine power is kept nearly constant at a speed when the en – route climb can start at a reduced power setting of the maximum climb rating. The psfc of the turboprop at takeoff is 0.5 lb/hr/shp, based on uninstalled power. There­fore, at SHPsls, the fuel-flow rate is 0.5 x 1,075 = 537.5 lb/hr. The intake airmass

Figure 13.5. Available turboprop thrust at take­off rating from a 1,100-SHP engine

flow at SHPsls is 0.011 x 1,075 = 11.83 lb/s (the specific power of 0.11 lb/s/SHP is provided in Section 10.2.2).

This type of turboprop engine is used in both civil and military aircraft design. The power settings are as in civil aircraft applications. Sizing of the Tucano class turboprop trainer aircraft requires a matched, installed engine TSLS = 4,000 lbs. Section 10.11.2 presents the generic, uninstalled, turboprop engine performance in nondimensional form. Section 10.11.4 works out the propeller thrust from the tur­boprop engine and establishes that the rated engine power of SHPSLS = 1,075 SHP (uninstalled) would develop as installed TSLS = 4,000 lbs.

Thrust computations from the turboprop SHP is repetitious work, as shown in Section 10.11.4. In this section, one computation each at the maximum climb rating and the maximum cruise rating, both at a 30,000-ft altitude for four speeds, are given in Tables 13.2 and 13.3, respectively.

Larger engines have a higher BPR. Current large operational turbofans have a BPR around 5 or 6 (new-generation turbofans have achieved a BPR > 8). These engines have performance characteristics slightly different than smaller engines – specifi­cally, the maximum climb rating has no break in thrust with altitude gain. Using Figures 10.48 through 10.50, the installed thrust and fuel-flow rates can be worked out as in previous sections.

13.3.2 Military Turbofan (Advanced Jet Trainer/CAS Role – Very Low BPR) – STD Day

This extended section of the book can be found on the Web at www. cambridge .org/Kundu and presents a typical military turbofan-engine installed performance (with and without reheat) at maximum rating suited to the classroom example of an AJT and a derivative in a CAS role. The installed performance is computed from the graph given in Subsection 10.11.4.

Table 13.2. 30,000-ft altitude (p = 0.00088 slug/ft3, a = 0.37) maximum climb

Depending on how the ECS is managed, installation loss typically varies from 6 to 8% of the uninstalled, sea-level static thrust. If required, the air-conditioning can be turned off for a brief period until the undercarriage is retracted. Using a 7% instal­lation loss at takeoff, Section 11.6 works out the matched installed TSLSjNSTALLED = 0.93 x 3,560 = 3,315 lbs per engine for the sized Bizjet. Figure 13.1 shows the installed engine thrust at the takeoff rating.

The fuel-flow rate is computed from the sfc of 0.498 lb/hr/lb at the sea-level, static condition (see Section 10.11.3). Using the uninstalled TsLs = 3,560 lbs per engine, the fuel-flow rate is 3,560 x 0.498 = 1,772 lbs per hour per engine. Fuel flow is kept nearly constant at takeoff up to the enroute climb segment, when the engine is throttled down to the maximum climb rating (computed in the following section).

Figure 13.1. Installed takeoff performance per engine (^BPR 3 to 4)

Maximum Climb Rating (Bizjet): Standard Day

Figure 10.46 shows the uninstalled maximum climb thrust in nondimensional form in terms of TSlS and fuel consumption (sfc) up to a 50,000-ft altitude for three Mach numbers. The installation loss during a climb is 6% of the uninstalled thrust. Using these graphs, the installed thrust and fuel-flow rates are plotted in Figure 13.2. This turbofan has a break in the fuel flow at a 5,000- to 10,000-ft altitude, depending on the flight Mach number, to keep the EGT within the limits, which results in a corresponding break in thrust generation (see Figure 13.2).

Equation 11.15 in Chapter 11 requires a factor k2 to be applied to the TSlS to obain the initial climb thrust. In the example, k2 is 1.5. Continuing with the coursework exercise, the uninstalled, initial climb thrust is 3,560/1.5 = 2,373 lbs per engine and the installed thrust becomes TSlS1nSTalled = 0.94 x 2,373 = 2,231 lbs per engine. Fuel flow at the initial climb is obtained from the sfc graph in Fig­ure 10.46b. For the initial climb, the sfc is 0.7 pound per hour per pound, which results in a fuel flow of 0.7 x 2,373 = 1,661 lbs/hr per engine. Equations for the climb performance are derived in Section 13.4.3 and the coursework example is verified in Section 13.5.2. Estimation of the payload range requires the full aircraft climb performance up to the cruise altitude.

= 600 –

C 400 =

Mach Number

Figure 13.3. Installed maximum cruise performance per engine (^<BPR 4)

Maximum Cruise Rating (Bizjet): Standard Day

Figure 10.47 shows the uninstalled maximum cruise thrust in nondimensional form and fuel consumption (sfc) from a 5,000- to 50,000-ft altitude for Mach numbers varying from 0.5 to 0.8. Figure 13.3 shows the installed-engine thrust at the maxi­mum cruise rating for the sized Bizjet.

The coursework example specified an initial maximum cruise speed (i. e., HSC) of Mach 0.74 at 41,000 ft. From Figure 10.47, that point gives the uninstalled ratio T/TSLS = 0.222 (TSLS/T = 4.5). This is the k in Section 11.3.3 that results in an uninstalled thrust of 3,560 x 0.222 = 790 lbs per engine. Considering a 4% installa­tion loss at cruise, the installed thrust of T = 0.96 x 790 = 758 lbs per engine. Sec­tion 13.5.3 verifies whether the thrust is adequate for an aircraft to reach the maxi­mum cruise speed. The fuel in Figure 13.4 (see Web at www. cambridge. org/Kundu) is 0.73 x 790 = 577 lb/hr per engine.

The discussion in this section generates the available installed thrust and fuel-flow graphs matched for the worked-out, sized-aircraft examples (see Chapter 11): a Bizjet and an AJT. In addition, the performance data for a 1,140-shp turboprop engine are provided for readers to work out the associated aircraft performance.

Because the given sfc graphs are based on uninstalled thrust, the fuel-flow rates are computed using uninstalled thrust. Installation loss at cruise is approximately half the percentage loss at takeoff.

13.3.1 Turbofan Engine (BPR < 4)

Figures 10.45 through 10.47 provide the typical uninstalled turbofan thrust in nondi­mensional form in terms of TsLs, along with the corresponding sfc for the Bizjet aircraft class. Section 11.6 establishes the requirement of an uninstalled matched TsLs_UNINSTALLED = 3,560 lb per engine. Worked out herein and summarized in Table 13.1 are examples of installed thrust and fuel flows for the three engine rat­ings. The data are sufficient for the example used in this book; intermediate values may be interpolated linearly.

Aircraft speed is a vital parameter in computing performance. It is measured using the difference between the total pressure, pt, and the static pressure, ps, expressed as (pt – ps). Static pressure is the ambient pressure in which an aircraft is flying. The value of (pt – ps) gives the dynamic head, which depends on the ambient air density, p. Unlike the ground speed of an automobile that is measured directly, an aircraft ground speed must be computed from (pt – ps); a pilot reads the gauge that is converted from (pt – ps). Following are various forms of aircraft air speed that engineers and pilots use. As shown, some computations are required – currently, onboard computers perform all computations:

Vi: The gauge reading as a pilot sees it in the flight deck; this is flight speed,

which is not the same as ground speed. The instrument includes stan­dard adiabatic compressible-flow corrections for high-subsonic flights at the sea-level standard day; however, it still requires other corrections.

VI: This is the indicated air speed (IAS). Manufactured instruments have

some built-in instrumental errors, Д Vi (typically minor but important considerations when an aircraft is close to stall speed). Manufacturers supply the error chart for each instrument. The instrument is calibrated to read the correct ground speed at the sea-level standard day with compressibility corrections. When corrected, the instrument reads the IAS as

Vi = IAS = Vi + Д Vi

VC: This is the calibrated air speed (CAS). Instrument manufacturers cali­

brate an uninstalled, bare instrument for sea-level conditions. Once it is installed on an aircraft and depending on where it is installed, the air­craft flow field distorts the instrument readings. Therefore, it requires position-error (Д Vp) corrections by the aircraft manufacturers:

Vc = CAS = Vi + Д Vp = Vi + ДVi + Д Vp

VeAS: This is the equivalent air speed (EAS). Air density p changes with alti­tude – it decreases because atmospheric pressure decreases with a gain in altitude. Therefore, at the same ground speed (also known as true air speed [TAS]), the IAS reads lower values at higher altitudes. The mathematical relationship between the TAS and the EAS, reflecting the density changes with altitude, can be derived as TAS = EAS/^/ст, where a is the density ratio (p/po) in terms of the sea-level value, p0. The constant EAS has a dynamic head invariant. For high-subsonic flights, it requires adiabatic compressibility corrections ^Vc) for the altitude changes:

Veas = EAS = Vc + ДVc. = Vi + Д Vi + Д Vp + Д Vc = TAS^a

TEAS: TAS is the aircraft ground speed. Compressibility corrections for posi­

tion errors are available; however, at this stage of design, the details can

be omitted without any loss of conceptual design work undertaken in this book. Supersonic flight requires further adjustments.

13.1 Overview

This chapter assesses whether the aircraft being configured, thus far, meets the FAR and customer requirements given in the form of specifications. Coursework fol­lows linearly from the mock market survey (see Chapter 2). Specification require­ments addressed in this chapter include aircraft performance to meet the (1) TOFL, (2) LFL, (3) initial rate of climb, (4) maximum speed at initial cruise (especially for civil aircraft design), and (5) payload range. Chapter 16 computes the aircraft DOC, which should follow the aircraft performance estimation.

Aircraft performance is a subject that aeronautical schools offer as a separate course. Therefore, to substantiate the FAR and customer requirements, this chap­ter addresses only what is required – that is, the related governing equations and computational examples associated with the five substantiation parameters listed previously. Substantiation of the payload range requires integrated performances of climb and descent that show fuel consumed, distance covered, and time taken for the flight segments. Integrated climb and descent performances are not specification requirements at this stage; therefore, their detailed computational examples are not provided. Instead, the final results in graphical form carry out the payload-range estimation. It is suggested that readers refer to appropriate textbooks for details on this topic. The turboprop example is not worked out but there is sufficient informa­tion to compute it similarly.

The remainder of the book after this chapter (except Chapter 16) presents infor­mation that aircraft designers should know and apply to their configurations. These topics may comprise the coursework of a second term following the finalized con­ceptual study in the first term. The discussion in Section 13.7 is useful to readers.

13.1.1 What Is to Be Learned?

This chapter covers the following topics:

Section 13.2: Preliminary information on aircraft performance

Section 13.3: Engine performance graphs

Section 13.4: Pertinent aircraft performance equations

Section 13.5: Performance equations to substantiate Bizjet aircraft capabilities

Section 13.6: Performance equations to substantiate AJT aircraft capabilities

Section 13.7: Discussion in summary form

13.1.2 Coursework Content

Readers perform the following steps for their design projects:

Step 1: Generate the appropriate engine performance graphs from the non­dimensional graphs provided in Chapter 10.

Step 2: Using the engine thrust thus obtained, compute the aircraft perfor­mances of the sized Bizjet and AJT as coursework exercises. (The instructor’s assistance may be required to compute integrated climb, descent, and specific-range performances.)

Step 3: If aircraft performance requirements are not met, then iterate the aircraft-configuration, sizing, and engine-matching exercises until they are. The spreadsheet method is helpful for the iterations.

13.2 Introduction

The final outcome of any design is to substantiate the performance it is intended to do. In the conceptual design phase, aircraft performance substantiation must be con­ducted mainly for those critical areas specified by the FAR and customer require­ments; a full aircraft performance estimation is conducted subsequently (it is beyond the scope of this book). All worked-out aircraft performance estimations (i. e., Bizjet and AJT) use the standard day. Non-ISA-day performance computations are calculated in the same way using non-ISA-day data.

The sizing exercises in Chapter 11 demonstrate a rapid-performance method to generate relationships between wing-loading (W/Sw) and thrust-loading (Tsls/W) to obtain the sizing point that simultaneously satisfies the requirements of the TOFL and LFL, initial rate of climb capability, and maximum speed at initial cruise. The aircraft-sizing point gives the installed, maximum sea-level takeoff static thrust, TSLSjnSTalled, of the matched engines. Chapter 10 presents the generic, uninstalled-engine performances of rubberized engines in nondimensional form, from which the installed-engine performances are obtained.

This chapter develops available engine performance in terms of installed thrust and fuel-flow rates at various speeds and altitudes at the power settings of takeoff, maximum climb, and maximum cruise ratings at standard day, matched for the sized aircraft under study. Applying the installed-engine data, the chapter continues with more accurate computations of aircraft performance to substantiate requirements of the TOFL and LFL, initial rate of climb, maximum speed at initial cruise, and payload range. At this point, it may be necessary to revise the aircraft configura­tion if performance capabilities are not met. If the aircraft performance indicates a shortfall (or an excess) in meeting the requirements, the design is iterated for improvement. In coursework, normally one iteration is sufficient.

Finally, at the end of the design stage, the aircraft should be flight-tested over the full flight envelope, including various safety issues, to demonstrate compliance.

It is clear that stability considerations are important in aircraft-design configura­tions. Although the related geometrical parameters are from statistical data of past designs and subsequently sized, this chapter provides a rationale for their role in the conceptual design stage. It also has been shown that to control inherent aircraft motions, feedback-control systems such as a stability augmentation system (SAS) (e. g., a yaw damper) and a control augmentation system (CAS) have been routinely deployed for some time. In this final section, the rationale continues with a discus­sion on how the feedback-control system has advanced to the latest technologies, such as FBW and fly-by-light (FBL), known collectively as ACT. Today, almost all types of larger aircraft incorporate some form of ACT.

The advantages of FBW are discussed in various sections of this book; the con­cept is not new. FBW is basically a feedback-control system based on the use of digital data. Figure 12.17 shows the control of one axis, which can be used for all three axes. Earlier SAS and CAS had mechanical linkage from the pilot to the controls; FBW does not have the direct linkage (hence, the name). It permits the transmission of several digital signal sources through one communications system, known as multiplexing. A microprocessor is in the loop that continuously processes air data (i. e., flight parameters) to keep an aircraft in a preferred motion with or without pilot commands. Aircraft-control laws – algorithms relating a pilot’s

command to the control-surface demand and aircraft motion, height, and speed, which involve equations of motion, aircraft coefficients, and stability parameters – are embedded in the computer to keep the aircraft within the permissible flight envelope. Under the command of a human pilot, the computer acts as a subservient flier. The computer continuously monitors aircraft behavior and acts accordingly, ensuring a level of safety that a human pilot cannot match.

Figure 12.17 is a schematic diagram of the FBW feedback arrangement for pitch control. The flight-control computer takes the pilot’s steering commands, which are compared to the commands necessary for aircraft stability to ensure safety and that control surfaces are activated accordingly. Air data are continually fed to the com­puters (i. e., speed, altitude, and attitude). Built into the computer are an aircraft’s limitations, which enables it to calculate the optimum control-surface movements. Steering commands are no longer linked mechanically from the cockpit to the con­trol surface but rather via electrical wiring. FBW flight-control systems seem to be the ideal technology to ensure safety and reduce a pilot’s workload.

Because analog point-to-point wire bundles are an inefficient and cumbersome means of interconnecting sensors, computers, actuators, indicators, and other equip­ment onboard a modern military aircraft, a serial digital multiplex data bus was developed. MIL-STD-1553 (in use since 1983) defines all aspects of the bus (i. e., a subsystem of electrical lines for communication, named after electrical bus bars); therefore, many groups working with the military have adopted it. The MIL-STD – 1553 multiplex data bus provides an integrated, centralized system control and a standard interface for all equipment connected to the bus. The bus concept pro­vides a means by which all traffic is available and can be accessed using a single connection for testing and interfacing with the system.

FBW reacts considerably faster than a conventional control system and does not encounter fatigue problems. A strong driver for incorporating FBW in military – aircraft design is the ability to operate at relaxed stability (even extending to a slightly unstable condition) used for rapid maneuver (increased agility) as a result of minimal stiffness in the system. It is difficult for a typical pilot to control an unsta­ble aircraft without assistance; a computer is needed and a regulator supplies the necessary stability. This system does not generate the natural stability of a conven­tional aircraft but automatically trims the aircraft to the preferred flight conditions. Progress in FBW systems depends to a great extent on the progress of onboard com­puter power.

An aircraft flying under relaxed stability using FBW does not have the same requirement for geometrical features to provide low stiffness and damping. Hence, stability and control-surface sizing are different than in a conventional design: They are smaller and, hence, lighter with less drag. This is what is meant by a CCV.

Stable designs already have a down-pitching force because of the position of the NP aft of the CG. Any balancing force must be generated by a larger down­ward lift of the H-tail. Again, this decreases the maximum possible lift and increases the trim drag. In an unstable layout (e. g., the CG moving aft), the elevator’s lift is directed upward to counterbalance the moment. In this way, the aircraft’s total lift is increased; the aircraft wing therefore can be designed to be smaller and lighter and still provide the same performance. There is another benefit from the use of an unstable design: In addition to the aircraft’s increased agility, there is a reduction in drag and weight.

The difference between a conventional and a CCV design is shown in Table 12.1 and Figure 12.18 (see [13]) for longitudinal stability. The wing area and MTOM of both designs were unchanged; the CCV design yielded a smaller H-tail area with a larger CG range. The directional stability exhibits similar gains with a smaller V-tail area, thereby further reducing the OEM and permitting a bigger payload.

In summary, FBW provides considerable advantages, as follows:

• a simple and flexible system architecture although its design is complex

• consistent handling

• automatic stabilization

• safe maneuvering to the envelope limits

• ability to integrate with a wide range of designs (e. g., slats and swing-wing)

• ability to integrate with engine control through FADEC and the thrust vector

• use of side stick controller – provides free space in the cockpit layout and weight-saving

• incorporates relaxed stability for rapid maneuver, yet uses smaller control sur­faces

• permits complex configurations for stealth aircraft, which may not be favor­able for aerodynamic considerations leading to unstable aircraft (e. g., the F117 Nighthawk)

• digital data-handling allows multiplexing, which saves weight

• overall weight reduction

• allows standardization

• failure detection

• fault isolation

• built-in tests and monitoring

Figure 12.18. Comparison between a conventional and a CCV design

FBW has been in use for nearly a half-century but the obvious advantages were kept secret for a long time for military reasons. Early development in the pioneer­ing stages progressed slowly with some mishaps. Nearly two decades later, civil avia­tion took bold steps and Aircraft Radio Inc. (ARINC) standards emerged to control FBW designs. The Airbus was the first aircraft to incorporate full FBW in a major project. The midsized A320 twin-jet aircraft is the first commercial transport aircraft to incorporate full FBW without manual override. The Habsheim (June 26, 1988) and the Bangalore (February 14,1990, near the author’s residence) disasters posed many questions; however, practically all midsized and larger transport aircraft cur­rently incorporate some form of FBW technology.

The FBW system architecture has built-in redundancies. During the 1980s, such systems had quadruple-redundant architecture in which each system works inde­pendently. Nowadays, with improved reliability, a triplex system (with voting and consolidation) dominates design. FBW can be applied to one, two, or all three axes of control; modern systems incorporate all three. MIL-STD-1553 specifies that all devices in the system must be connected to a redundant pair of buses, which pro­vides a second path for bus traffic if one bus is damaged. Signals are allowed to appear only on one of the two buses at a time. If a message cannot be completed on one bus, the bus controller may switch to the other bus. In some applications, more than one bus may be implemented on a given aircraft.

To avoid electromagnetic interference, the use of fiber optics for signaling using light was developed recently. Aptly, it is called FBL and is guided by MIL-STD – 1773.

In summary, FBW and FBL designs offer weight reduction with a smaller wing and empennage, fewer control surfaces, less cabling, and the elimination of mechanical linkages. As a consequence, drag is reduced. In addition, FBW and FBL designs provide enhanced safety and reliability as well as improved failure detection.