Category AIRCRAF DESIGN

MFD started as a display on a cathode ray tube (CRT) but has advanced to a liq­uid crystal display (LCD), which is a lighter and clearer technology. All relevant data for pilot use (e. g., air, engine, and navigational data) are displayed simul­taneously on the screen. To reduce clutter, the displays are divided into primary and secondary displays. Separate-system displays are accommodated in one or two

EFISs: the primary air-data system display (SD), and the navigational display (ND); each type of system has several pages and each display screen can be changed for specific information. Figure 15.15 shows typical EFIS displays. EFIS/MFD/ND/SD have many pages that can be flipped to as desired, including pages for the engine, cruise, flight-control, fuel, electrical, avionics, oxygen, air-bleed, air-conditioning, cabin-pressurization, hydraulics, undercarriage, doors, and the APU (military air­craft have weapons-management pages).

The primary flight display (PFD) consists of air-data systems, including aircraft speed, altitude, attitude, aircraft reference, and ambient conditions. The secondary system consists of the ND, which provides directional bearings (i. e., GPS and inertial system), flight plan, route information, weather information, airport information, and so forth. For pilot facility, each type has some duplication. In a separate panel, the SD shows the engine data and all other system data, including those required for the ECS. EFISs have removed the clutter of analog dials, one for each type of data. In some designs, the engine display (ED) is shown separately. Forward­looking weather radar can have the ND or a separate display unit.

Initially, flight decks also had basic analog gauges showing air data as redundan­cies in case the EFIS failed. Currently, with vastly improved reliability in the EFIS, older analog gauges are gradually being removed.

The aircraft flight deck is a better term than the older usage of the word cockpit, which originated in ship design in the sixteenth century; it was similar to men working in a confined area under stress, like cocks that were forced to fight in a pit for sport. Crew station is another term meaning the same as a workplace for operators of any type of vehicle. To standardize terminology, this book uses flight deck, intended specifically for aircraft. The flight deck serves as a human-machine interface by providing (1) an outside reference of topography through the cabin win­dows, (2) onboard instruments to measure flight parameters, (3) control facilities to operate an aircraft safely for the mission role, and (4) management of aircraft sys­tems (e. g., the internal environment). Future designs with advanced displays could result in a visually closed flight deck (i. e., a TV replacing the windows). The front – fuselage shape can be influenced by the flight-deck design. Transport aircraft have two pilots sitting next to one another at a pitch of about 1.2 meters in smaller air­craft to 1.4 meters in larger aircraft. Understanding the flight-deck arrangements also provides a sense of the equipment requirements that result in a measure of the associated weights involved. The space and adequacy of vision polar, which estab­lishes the window-size requirements, also can be better understood.

Both civil and military aircraft pilots have the following common functions:

• mission management (planning, checks, takeoff, climb cruise, descent, and landing)

• flight-path control

• systems management

• communication*

• navigation*

• routine postflight debriefing

• emergency action when required (drills differ between civil and military aircraft)*

*Civil aircraft pilots are assisted by ground control (i. e., communication and naviga­tion), whereas in a critical situation, combat pilots must manage the aircraft them­selves – which is a significant difference. Both situations may require taking emer­gency actions, but for a combat pilot, this could be drastic in nature (i. e., ejection; see Section 15.10). In addition, military aircraft pilots have an intense workload, as follows:

• mission planning (e. g., Lo/High combination; see Chapter 13); this is required for mission management (preflight briefing may change if the situation demands)

• target acquisition

• weapons management and delivery

• defensive measures and maneuvers

• counterthreats; use of tactics

• management of situation when hit

• in-flight refueling, where applicable

• detailed postflight briefing in special situations

The military aircraft flight deck is under more stringent design requirements. The civil aircraft flight-deck design is in the wake of military standards and the provi­sion of space is less constrained. This is why the military aircraft flight deck is dis­cussed first (see Figure 15.16). An aircraft flight-deck design has changed dramati­cally since the early analog-dial displays (i. e., four-engine aircraft gauges now fill the front panel; see Figure 15.17) to modern microprocessor-based data management in an integrated, all-glass, multifunctional display (MFD), which is also known as an electronic flight information system (EFIS).

Emergency situations (e. g., fire hazard and ditching on water or land) require a fast exit from the aircraft cabin to safety. The FAA initially imposed a 120-s egress time but, in 1967, changed it to a maximum of 90 s. This was feasible through advances made in slide and chute technology. To obtain an airworthiness certi­fication, an aircraft manufacturer must demonstrate that complete egress is pos­sible within 90 s by conducting simulated tests. The EASA has similar require­ments.

FAR Part 25, Sections 25.783 and 25.807, give requirements for the main cabin doors and emergency exit doors, respectively. Several types of emergency exit doors are listed in Table 15.7 (in inches); all are rectangular in shape with a corner radius. The sizes are a minimum size and designers can make them larger. Oversized doors need not be rectangular as long as the minimum rectangular size is inscribed.

All doors except Type III (i. e., an inside step up of 20 inches and an outside step down of 27 inches) and Type IV (i. e., an inside step up of 29 inches and an outside step down of 36 inches) are from the floor level. If a Type II door is located over the wing, it can have an inside step up of 10 inches and an outside step down

of 17 inches. Emergency doors are placed at both sides of the aircraft and do not need to be diametrically opposite; however, they should be uniformly distributed (i. e., no more than 60 ft apart) and easily accessible for evenly distributing loading passengers when required. The safety drill by the cabin crew is an important aspect in saving lives, and all passengers should listen to the demonstration regardless of how frequently one flies. There are differences among door types.

An aircraft should have at least one easily accessible external main door. The combination of main and emergency doors is at the discretion of the manufac­turer, which must demonstrate a simulated evacuation within the stipulated time. The fuselage length also determines the number of emergency doors because they should not be spaced more than 60 ft apart. Table 15.8 lists the minimum number of emergency doors; it is recommended that more than the minimum be provided. Types A, B, and C also can be used and they are deployed in larger aircraft.

There may be other types of doors such as a door at the tail cone and ventral doors, the dimensions of which are listed in Table 15.9. Flight-crew emergency-exit doors are provided separately in the flight deck.

When the door level is high above the ground, inflatable escape slides and chutes are provided, as shown in Figure 15.14. In an emergency situation in which stairs may not be available (or there may not be time to wait for them to arrive), inflatable chutes are used for passenger evacuation within the specified time. The slides and chutes also serve as rafts with floating attachments.

As aircraft size increases, the technological demand to facilitate quick egress becomes a more challenging task. In March 2006, the Airbus 380 demonstrated that 850 passengers could be evacuated in 80 sec (although with minor injuries). How­ever, a typical Airbus 380 passenger load is fewer than 650 passengers with a mixed – class arrangement. The Airbus has sixteen exits but was successful in the evacuation using only eight doors (i. e., half remained closed).

Figure 15.14. Inflatable escape chute and slide

Coursework Exercise

There is a coursework exercise in this chapter. The configuration developed in Chapter 6 is to be reverified. The Bizjet must have the following features:

It is best for all doors to have Type III standards for component commonality, which reduces production costs.

Creating just the aircraft shell, satisfying only aerodynamic needs, has consequences during manufacture. It is simple to create the drawings but not as easy to produce the hardware. During the conceptual study phase, it is routine procedure in the indus­try to obtain the valued opinion of production engineers in an IPPD environment. Compromises may be made in shaping an aircraft if doing so facilitates manufac­turability, which in turn saves cost – more so in the commercial aircraft business, where operational economic gains are more important than in pure aerodynamics.

Manufacturing philosophy is associated with the choice of materials, machining routine, forming, fabrication, and assembly-tool (i. e., jigs and fixtures) concepts (see Chapter 17). Typically, the aim is to shape components as simply as possible with fewer parts and faster assembly time. Attention is given to minimizing complex 3D curvatures; applying more circular shapes than complex, convoluted curves; main­taining commonality of geometry; and providing accessibility for maintenance.

Therefore, it is suggested that a second-term project be assigned to introduce the structural philosophy in harmony with the manufacturing philosophy, beginning with internal structural arrangements in simple line diagrams, as shown for an aft – fuselage in Figure 15.13a. Similar line drawings for the wing and empennage are not shown herein. The advantages of using CAD are discussed in previous chapters of this book, which are apparent, as shown for a typical military aircraft fuselage in Figure 15.13b. If a basic aircraft configuration is created in CAD, then the external aircraft contour lines guide the shape of the internal arrangement, with the added benefit of being able to examine accessibility and production complexity to establish manufacturing philosophy.

The strategy is to lay out the main internal structural arrangements, such as the position of the ribs, spars, longerons, bulkheads, wing box-fuselage, flap – empennage attachment, engine attachment, and fuel tank. At this stage, it is not detailed component design. In the next phase (i. e., Phase 2, project definition), line diagrams are developed into shapes after stressing to consolidate the manufactur­ing philosophy and, if necessary, to prepare for the bidding process to subcontract work. During Phase 3 (i. e., detailed design), the parts are developed into detailed production drawings ready for manufacture. The use of CAD avoids duplication in generating components and the subassemblies outline. CAD is capable of making the procedures paperless. Reference [5] provides a good description and analyses of aircraft structural design.

Aluminum alloys continue to be the most prevalent material in aircraft structure. Table 15.6 is a conservative presentation of typical percentages of composite use for the coursework project. The table reflects the typical current practice, although there are some newer designs that perform better than what is listed.

The introductory coursework exercise may use the following strategy. The weight equations provided in Chapter 8 are valid.

Civil Aircraft Design

For civil aircraft, the design is all-metal construction. If nonmetals (i. e., compos­ites) are introduced, they may be limited to secondary and tertiary structures up to only approximately 25% of the OEW. Typical nonmetal structures include the floorboards, control surfaces, complex fairings, and empennage.

Military Aircraft Design

For military aircraft design, the same philosophy about nonmetal (i. e., composites) components is maintained, with weights increased to 40% of the OEW. Typical non­metal structures include the floorboards, control surfaces, complex fairings (e. g., intake ducting and wing-body junction), and empennage.

Material selection for any engineering product depends on its function, shape, man­ufacturability (i. e., process), and cost. For aircraft applications, there is the addi­tional consideration of weight. Within the material classes, there are subclasses of alloys: composites that offer appropriate properties to suit a product – a large variety is available with ever-increasing newer types. In the conceptual design phase, engi­neers must screen and rank the types of materials that suit the requirements, listing the limitations and constraints involved and whether a change to another type is

Several options are available for appropriate materials to make the best compro­mise. Thus, aircraft-weight estimation is more complex, and engineers must identify and compute numerous parts to estimate component weights before an aircraft is built; CAD 3D modeling helps.

Choice of material affects aircraft weight and cost. The semi-empirical relation for weight estimation in Chapter 8 considers all-metal construction and describes how to adjust the prediction if some parts are made of a lighter material. For a rapid method, the OEW may be factored accordingly – only the structural weight is affected; the remainder is unchanged. Composites may be used in secondary and tertiary structures, where loads are low and failure does not result in catastrophe.

In general, for the same Young’s Modulus, metals have higher density. How­ever, when the strength-to-weight ratio (i. e., specific strength) is considered, then composites overtake metals; that is, engineers can obtain the same strength with lighter components even when the higher FS erodes the weight savings. Metals demonstrate a better Young’s Modulus for the same strength. Metals also show better fracture toughness for the same Young’s Modulus. Another important com­parison is the relative cost per unit volume versus the Young’s Modulus when metal alloys are less costly.

Under load (i. e., stress), all materials deform (i. e., strain) – some more than others – but they can recover their original shape when the load is removed, provided that the application is within a specific limit. Beyond this load level, materials do not recover to their original shape. See [4], [5], and [6] for details on stress and strain.

Stress is the applied force per unit area of a material. It is termed as tensile or compressive stress when the force is acting normal to an area and shear stress when it is acting tangentially. The associated deformation per unit length or area is the normal or shear strain, respectively. How a material is prepared affects the

Stabilizer tips

Strut fairings

Cowl components

Wing to body fairing

T. E. flap linkage fairing

Outboard

ailerons

Strain

(a) Aluminum alloy

Figure 15.12. Material stress-strain relationship

characteristics of the stress-strain relationship. The nature of alloys, crystal forma­tion, heat treatment, and cooling affects a materials characteristics.

A typical stress-strain characteristic of an aluminum alloy is shown in Fig­ure 15.12. The figure shows that initially, the stress-strain relationship behaves lin­early according to Hooke’s Law, which represents the elastic property of a material. Within the elastic limit, the material strength (i. e., how much stress it can bear) and stiffness (i. e., how much deformation occurs) are the two main properties con­sidered by designers in choosing materials; of course, the cost, weight, and other properties are also factors to consider. The maximum point within which linearity holds is the yield point. Past the yield point, permanent deformation occurs: The material behaves like plastic and the slope is no longer linear. The highest point in the stress-strain graph is known as the ultimate strength, beyond which the compo­nent continues to deform and results in a rupture that is a catastrophic failure. The linear portion gives the following:

stress/strain = constant = Young’s Modulus (the slope of the graph)

Sometimes raw material is supplied with a small amount of prestretching (i. e., strain hardening) and a permanent deformation set in which the yield point is higher. Typically, some aluminum sheet metals are supplied with 0.2% built-in prestretched strain (see Figure 15.12a). However, with prestretching, the ultimate strength is unchanged. Figure 15.12b compares various types of typical aircraft materials. A steeper slope indicates higher stiffness, which often has a higher elas­tic limit. Brittle materials rupture abruptly with minimal strain buildup; ductile materials exhibit significant strain buildup before rupturing, thereby warning of an imminent failure. Rubber-like materials do not have a linear stress-strain relation­ship. The pertinent properties associated with materials follow (some are shown in Figure 15.12): [28]

• hardness: a measure of strength

• resilience: a measure of energy stored in an elastic manner; that is, the strain is restored when the stress is relieved

• toughness (fracture toughness): a measure of resistance to crack propagation

• creep resistance: a slow deformation with time under load; strain can increase without applying much stress

• wearability: a measure of surface degradation mainly under exposure (e. g., cor­rosion)

• fatigue quality: set up with alternate cycling of applied load; a good material dissipates vibration energy for a number of cycles

• ability to hold strength at elevated temperature

The limit load, ultimate load, and factor of safety (FS) associated with material are described in Section 5.5.2. Limit load is up to the point where there is no perma­nent deformation under load. Certifying agencies stipulate strict control on aircraft structural integrity. For unpredictability (e. g., under a gust load or material defect), an FS is incorporated to accept an ultimate load when some deformation is allowed but is still below the ultimate strength. For metals, the FS = 1.5; that is, a 50% increase from the limit load is allowed. The properties of composite materials have reduced values of the stress level to allow for damage tolerance and environmental issues and to maintain an FS of 1.5 (see Section 5.6). The manufacturing process also determines the allowable stress level. These considerations can penalize part of the weight-saving associated with using lighter materials.

The strength and other properties vary among materials. Table 15.4 lists impor­tant materials used in the aircraft industry (only typical values are given). Wood has many variations and is not used much anymore.

With an increase in temperature, material properties degrade. Special alloys of steel and titanium retain better strength at elevated temperatures. Components experiencing a hot temperature have titanium and stainless-steel alloys that are available in many variations. In the quest to find still-better materials, nickel, beryl­lium, magnesium, and lithium alloys have been produced. The more exotic the nature of an alloy, the more costly it is. Typically, an aluminum-lithium alloy is three to four times more expensive than duralumin (in 2005). Aluminum alloys are still the dominant material used in the aircraft industry. The variety of alu­minum alloys indicates a wide range available for specific uses. Various types of alu­minum alloys are designated (i. e., classified) with a numbering system, as shown in Table 15.5.

Aircraft that defy gravity necessarily must be weight-efficient, thereby forcing designers to choose lighter materials or – more precisely – those materials that give a better strength-to-weight ratio. Also implied are the questions of cost of raw mate­rials, cost of fabrication, and stability during use. This section helps readers under­stand that choosing the appropriate materials is an involved topic and therefore is an integral part of the study during the conceptual design phase. Aircraft weight and cost are affected by the choice of materials and, therefore, aircraft performance and economy. The success or failure of a new aircraft design depends largely on the choice of appropriate materials, especially when the number of those available is increasing.

In the early days of aviation, the only choice was to use an all-wood construc­tion or a fabric cover to wrap around a wooden airframe to serve as an aerody­namic surface. Being anisotropic and without enough resistance to impact, wood properties have limitation. At that time, the available metals were heavy and the lighter ones were soft and corrosive. Today, wood is no longer used except in the homebuilt-aircraft category, primarily because wood is the easiest material with which to work. Moreover, the ethical question of forest conservation discourages the use of wood.

In the 1920s, the combination of progress in engines and in aerodynamic tech­nologies allowed aircraft speed to exceed 200 mph, which required better mate­rials. Technology changed in the 1930s when Durener Metallwerke of Germany introduced duralumin, an alloy of aluminum, with a higher strength-to-weight ratio, improved anticorrosion properties, and isotropic properties. The company followed with a variety of alloys for specific manufacturability, damage tolerance, crack prop­agation, and anticorrosive properties in the form of clad-sheets, rolled bars, ingots, and so forth. The introduction of metal also resulted in a new dimension to manufac­turing philosophy. Progress in structures, aerodynamics, and engines paved the way for substantial gains in speed, altitude, and maneuverability performance. These improvements were seen primarily in the World War II designs, such as the Super­marine Spitfire, the North American P-51, the Focke Wolfe 190, and the Mitsubishi Jeero-Sen.

The last three decades have seen the appearance and increasing use of non­metals, such as fiberglass/epoxy, kevlar/epoxy, and graphite/epoxy, which are com­posite materials constructed in layers of fabric and resin. Composites have bet­ter strength-to-weight ratios compared to aluminum alloys, but they also have anisotropic properties. Because they are shaped in moulds during the fabrication of parts, difficult curvy 3D shapes can be produced relatively easily. The near future will see more variety of composite materials embedded with metal to obtain the best of both. The Bombardier CSeries, Airbus 380, Boeing 787, and Airbus 350 are examples of how extensively composite materials are used. The technology of com­posite materials is evolving at a fast rate and there will be more variety in composite materials with better properties and capabilities at a lower cost.

Typically, composites may be used in secondary and tertiary structures in which loads are low and any failure does not result in catastrophe. Figure 15.11 shows the composite materials in a Boeing 767 aircraft. As the technology progresses, more composites will appear in aircraft moving into primary load-bearing structures.

Table 15.3 compares the extent of increase in composites from an older B747 (1960s) to the relatively newer design of the B777 (1990s). The latest B787 and A350 have considerably higher percentages of composites.

Composite materials are incorporated increasingly in percentage by weight. A few smaller aircraft are made of all composite materials but the FAR Part 23/25 cer­tification procedure is more cumbersome than for metal construction. It was difficult to obtain airworthiness certification for early all-composite aircraft because there were insufficient data to substantiate the claims. Military certification standards for aircraft structures are different.

Table 15.3. Percentage mass of types of material used in the aircraft structure

The newer military aircraft designs use expensive, exotic materials (e. g., aluminum-lithium alloy and boron alloy) that have yet to prove their cost – effectiveness in commercial aircraft. More than half of the Eurofighter’s structural mass is constructed of various types of composite materials; a fifth is made of the aluminum-lithium alloy.

Currently, the civil aviation sector burns about 12% of the fossil fuels consumed by the worldwide transportation industry. It is responsible for an approximate 3% annual addition to greenhouse gases and pollutant oxide gases. The environmental debate has become intense on issues such as climate change and depletion of the ozone layer, leading to the debate on long-term effects of global pollution. Smog consists of nitrogen oxide, which affects the pulmonary and respiratory health of humans. The success of the automobile industry in controlling engine emissions is evident by dramatic improvements achieved in many cities.

The U. S. Environment Protection Agency (EPA) recognized these problems decades ago. In the 1980s, the need for government agencies to tackle the engine emissions issue was emphasized. The early 1990s brought a formal declaration (i. e., the Kyoto agreement) to limit pollution (specifically around airports). Currently, there are no regulations for an aircraft’s cruise segment. In the United States, FAR Part 35, and internationally, the ICAO (i. e., Annexure 16, Volume II), outline the emissions requirements. EPA has worked closely with both the FAA and the ICAO to standardize the requirements. Although military aircraft emissions standards are exempt, they are increasingly being scrutinized for MILSPECS standards. Emissions are measured by an emission index (EI).

Combustion of air (i. e., oxygen plus nitrogen) and fuel (i. e., hydrocarbon plus a small amount of sulphur) ideally produces carbon dioxide (i. e., CO2), water (i. e., H2O), residual oxygen (i. e., O2), and traces of sulphur particles. In practice, the combustion product consists of all of these plus an undesirable amount of pollu­tants, such as carbon monoxide (i. e., CO, which is toxic), unburned hydrocarbons (UHC), carbon soot (i. e., smoke, which affects visibility), oxide of sulphur, and var­ious oxides of nitrogen (i. e., NOX, which affects the ozone concentration). The reg­ulations aim to reduce the level of undesirable pollutants by improving combustion technology. The sustainability of air travel and growth of the industry depend on how technology keeps up with the demands for human-health preservation.

Lower and slower flying reduces the EI; however, this conflicts with the market demand for flying higher and faster. Designers must make compromises. Reduction of the EI is the obligation of the engine manufacturer; therefore, details of the air­worthiness EI requirements are not provided herein (refer to the respective FAA and ICAO publications). Aircraft designers must depend on engine designers to supply certified engines that comply with regulatory standards.

Noise is produced by pressure pulses in air generated from any vibrating source. The pulsating energy is transmitted through the air and is heard within the audible frequency range (i. e., 20 to 20,000 Hz). The intensity and frequency of pulsation determine the physical limits of human tolerance. In certain conditions, acoustic (i. e., noise) vibrations can affect an aircraft structure. Noise is perceived as environ­mental pollution.

The intensity of sound energy can be measured by the sound pressure level (SPL); the threshold of hearing value is 20p, Pa. The response of human hearing can be approximated by a logarithmic scale. The advantage of using a logarithmic scale for noise measurement is to compress the SPL range extending to well over a million times. The unit of noise measurement is a decibel, abbreviated to dB, and is based on a logarithmic scale. One “Bel” is a tenfold increase in the SPL; that is, 1 Bel = logi010, 2 Bel = logi0100, and so on. A reading of 0.1 Bel is a dB, which is antilogi00.1 = 1.258 times the increase in the SPL (i. e., intensity). A twofold increase in the SPL is log102 = 0.301 Bel, or 3.01 dB.

Technology required a meaningful scale suitable to human hearing. The units of noise continued to progress in line with technology demands. First was the “A-weighted” scale, expressed in dB(A), that could be read directly from cali­brated instruments (i. e., sound meters). Noise is more a matter of human reaction to hearing than just a mechanical measurement of a physical property. Therefore, it was believed that human annoyance is a better measure than mere loudness. This

dB(A) CHART

Jet takeoff (25 m) (eardrum rupture)

Aircraft carrier deck Jet takeoff (100 m)

Live rock music, chain saw (threshold ofpain) Riveting, car hom (1 m)

Jet takeoff (305 m), lawn mower Busy urban street, food blender Dishwasher, factory, freight train (15 m) Vacuum cleaner, freeway traffic (15 m) Conversation in restaurant, office Quiet suburb, conversation at home Library

Quiet rural area Whisper, rustling leaves Breathing

Threshold of hearing

resulted in the “perceived-noise” scale expressed in PNdB, which was labeled as the associated “perceived noise level (PNL),” shown in Figure 15.1 from various origins.

Aircraft in motion presented a special situation with the duration of noise ema­nating from an approaching aircraft passing overhead and continuing to radiate rearward after passing. Therefore, for aircraft applications, it was necessary to intro­duce a time-averaged noise – that is, the effective perceived noise level (EPNL), expressed in EPNdB.

In the 1960s, litigation from damages caused by aircraft noise caused the gov­ernment regulatory agencies to reduce noise and impose EPNdB limits for various aircraft classes. Many airports have a nighttime curfew for noise abatement and control, with additional fees being charged for using the airfield at night. Through research and engineering, significant noise reduction has been achieved despite the increase in engine sizes that produces several times more thrust.

The United States was first to impose noise certification standards for aircraft operating within that country. The U. S. airworthiness requirements on noise are governed by FAR Part 36. An aircraft MTOM of more than 12,500 lb must comply with FAR Part 36. The procedure was immediately followed by the international agency governed by ICAO (see Annexure 16, Volume I). The differences between the two standards are minor, and there has been an attempt to combine the two into one uniform standard. Readers may refer to FAR Part 35 and ICAO Annexure 16 for further details.

Because existing larger aircraft caused the noise problem, the FAA introduced regulations for its abatement in stages; older aircraft required modifications within

Figure 15.2. Noise measurement points at takeoff and landing

a specified period to remain in operation. In 1977, the FAA introduced noise-level standards in three tiers, as follows:

Stage I: Intended for older aircraft already flying and soon to be phased out

(e. g., the B707 and DC8). These are the noisiest aircraft but least penalized because they are soon to be grounded.

Stage II: Intended for recently manufactured aircraft that have a longer life­

span (e. g., the B737 and DC9). These aircraft are noisy but must be modified to a quieter standard than Stage I. If they are to con­tinue operating, then further modifications are necessary to bring the noise level to the Stage III standard.

Stage III: Intended for new designs with the quietest standards.

Stage IV: Further increased stringency was applied for new aircraft certifica­tion during 2006.

ICAO standards are in Annexure 16, Volume I, in Chapters 2 through 10, with each chapter addressing different aircraft classes. This book is concerned with Chap­ters 3 and 10, which are basically intended for new aircraft (i. e., first flight of a jet aircraft after October 6, 1977, and a propeller-driven aircraft after November 17, 1988).

To certify an aircraft’s airworthiness, there are three measuring points in an airport vicinity to ensure that the neighborhood is within the specified noise limits. Figure 15.2 shows the distances involved in locating the measuring points, which are as follows:

1. Takeoff reference point: 6,500 m (3.5 nm) from the brake release (i. e., starting) point and at an altitude given in Table 15.1.

2. Approach reference point: 2,000 m (1.08 nm) before the touchdown point, which should be within 300 m of the runway threshold line and maintained at least at a 3-deg glide slope with an aircraft at least at a 120-m altitude.

3. Sideline reference point: 450 m (0.25 nm) from the runway centerline. At the sideline, several measuring points are located along the runway. It is measured on both sides of the runway.

Make linear interpolations for in-between aircraft masses. For takeoff, make linear interpolations for in-between mass.

The arithmetic sum of noise levels at the three noise measuring position is referred to as the “cumulative noise level”; and the difference between this level and the arithmetic sum of the noise limits allowed at each measuring point is referred to as the “cumulative noise level margin.”

The maximum noise requirements in EPNdB from ICAO, Annexure 16, Vol­ume I, Chapter 3, are listed in Table 15.1 and plotted in Figure 15.3.

Approach

This is for any number of engines

MTOM (kg) <35,000

EPNdB limit 98

Sideline

This is for any number of engines (use linear interpolations for in-between masses).

MTOM (kg) <35,000 >400,000

EPNdB limit 94 103

A typical footprint of the noise profile around a runway is shown in Figure 15.4. The engine cutback area is shown with the reestablished rated thrust for an enroute climb. Residential developments should avoid the noise-footprint areas.

Stage IV requirements for new type designs from January 1,2006 are as follows:

• a cumulative margin of 10 EPNDB relative to Stage 3

• a minimum sum of 2 EPNDB at any 2 conditions

• no trades allowed

The airframe also produces a significant amount of noise, especially when an aircraft is in a “dirty” configuration (e. g., flaps, slats, and undercarriage deployed). Figure 15.5 shows the sources of noise emanating from the airframe. The entire wetted surface of an aircraft – more so by the flow interference at the junction of two bodies (e. g., at the wing-body junction) – produces some degree of noise based on the structure of the turbulent flow causing pressure pulses that are audible to the human ear. The noise is aggravated when the undercarriage, flaps, and slats are deployed, creating a considerable vortex flow and unsteady aerodynamics; the fluctuation frequencies appear as noise. In the conceptual design phase, care must be taken to minimize gaps, provide fillets at the two-body junction, make stream­lined struts, and so forth. Noise increases as speed increases. Care must be taken to eliminate acoustic fatigue in structures and to design them to be damage-tolerant; material selection is important.

Figure 15.5. Typical sources of noise emanating from an airframe

Figure 15.6. Relative noise distributions from various aircraft and engine sources

Typical noise levels from various sources are shown in Figure 15.6 at both take­off and landing. Aircraft engines contribute the most noise, which is reduced at land­ing when the engine power is set low and the jet efflux noise is reduced substantially. There is more noise emanating from the airframe at landing due to higher flap and slat settings, and the aircraft altitude is lower at the measuring point than at the take­off measuring point. Because the addition of noise level is in a logarithmic scale, the total noise contribution during takeoff and landing is almost at the same level.

The power plant constitutes the nacelle and is the main sources of noise at take­off when an aircraft is running at maximum power. All of the gas turbine compo­nents generate noise: fan blades, compressor blades, combustion chamber walls, and turbine blades. With an increase in the BPR, the noise level decreases because a low exhaust velocity reduces the shearing action with ambient air. The difference in noise between an AB turbojet and a high-BPR turbofan can be as high as 30 to 40 EPNdB. Figure 15.7 shows that in subsonic-flight speed, noise radiation moves ahead of an aircraft.

To reduce noise levels, engine and nacelle designers must address the sources of noise, as shown in Table 15.2. The goal is to minimize radiated and vibrational noise. Candidate areas in engine design are the fan, compressor, and turbine-blade; gaps in rotating components; and, to an extent, the combustion chamber. Engines are bought-out items for aircraft manufacturers, which must make compromises between engine cost and engine performance in selecting what is available on the market. Aircraft and engine designers communicate constantly to make the best choice without compromising safety.

Figure 15.7. In-flight turbofan noise – radiation profile

Table 15.2. Nacelle and turbofan technological challenges to reduce noise

Fan/compressor and turbine Burner

Internal liner – intake

• absorbs fan noise Internal liner – casing/fan duct

• insulates compressor noise

• absorbs burner noise Internal liner – jet pipe

• absorbs turbine noise

• absorbs burner noise

• mixes hot and cold flows

• improves exhaust-flow mixing

Figure 15.8 shows the positions of noise-suppression liners placed in various areas and the jet-pipe-flow mixing arrangements for noise abatement. Exhaust-noise suppression also is achieved by using a fluted duct (which increases the mixing area) at the exit plane. Many types of liners are available on the market and there is room for improvement in liner technology. Primarily, there are two types of lin­ers: reactive and resistive. The reactive liners have different sizes of perforations to react with matched frequency range of noise and absorb. The resistive type of liner is a noise insulator in layers with screens. The most common type of acoustic liner comprises a combination of both types. It has resistive facing sheet covering a honeycomb structure between the insulator screens with cell sizes matched to the frequency range where noise attenuation is requuired. Nacelle certification is the responsibility of the aircraft manufacturer, even when it is subcontracted to a third party, because it is covered by FAR Part 25 requirements, rather than FAR Part 33, which are for the engine.

Propeller-driven aircraft must consider noise emanating (i. e., radiation and reflection) from the propellers. Here, the noise-reflection pattern depends on the direction of propeller rotation, as shown in Figure 15.9. The spread of reflected noise also depends on the propeller position relative to the wing and the fuselage.

Inside the aircraft cabin, noise comes from the ECS and must be maintained at the minimum level. These problems are addressed by specialists. Cabin-interior design considerations are addressed in Phase 2 of a project.

15.3.1 Summary

At this stage of study, design considerations for noise reduction do not substan­tially affect the aircraft external configuration other than using proper filleting at

two-body junctions, streamlining the projected structure, minimizing gaps, and so forth. The finalized aircraft configuration – as obtained in Chapter 6 and sized in Chapter 11 – remains unaffected because the aircraft external geometry is assumed to have accounted for these considerations. The choice of materials (e. g., nacelle lin­ers, cabin insulators, and fatigue-resistant material) can affect aircraft mass. Engine – noise abatement is generally the responsibility of engine designers.

The advancement of CFD capabilities in predicting noise has resulted in good judgments for improving design. Substantiation of the CFD results requires testing.

In the near future (i. e., gradually evolving in about two decades), remarkable improvement in noise abatement may be achieved using a multidisciplinary design approach, taking the benefits from various engineering considerations leading to a BWB shape. Cambridge University and the Massachusetts Institute of Technol­ogy have undertaken feasibility studies that show a concept configuration in Fig­ure 15.10 for an Airbus 320 class of subsonic-jet commercial transport aircraft. The engineers predict that the aircraft will be 25 dBs quieter than current designs – so quiet as to name it “silent aircraft.” The shaping of the aircraft is not based solely on noise reduction; it is also driven by general aerodynamic considerations (e. g., drag reduction and handling qualities). Noise reduction results from the aircraft body shielding the intake noise, minimizing two-body junctions by blending the wing and the fuselage, and eliminating the empennage. Of course, reduction in the engine noise is a significant part of the exercise. However, to bring the research to a mar­ketable product will take time, but the author believes it will come in many sizes; heavy-lift cargo aircraft are good candidates.