Aircraft Landing Gear Design: Principles and Practices Norman S. Currey L o c k h e e d A e r o n a u t i c a l Systems C o m p a n y Marietta, Georgia AIAA. Aircraft Landing Gear Design: Principles and Practices Download the Full PDF from the initial concepts of landing gear design through final detail design. Download as PDF, TXT or read online from Scribd. Flag for . Aircraft Landing Gear Design: Principles and Practices Norman S. Currey Google Livros.
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Aircraft Landing Gear Design (Principles and Practices) - Norman S. Currey. downloadd from American Institute of Aeronautics and Astronautics Aircraft Landing Gear Design: Principles and Practices Norman S. Currey. PDF | A computer aided graphical synthesis was undertaken to Also of importance in the design of aircraft landing gear is the structure.
Site Tools Sign up for e-alerts What is this? Principles and Practices Norman S. Currey eISBN: PDF not available for download at this time. This is a comprehensive text that will lead students and engineers from the initial concepts of landing gear design through final detail design.
The book provides a vital link in landing gear design technology from historical practices to modern design trends, and it considers the necessary airfield interface with landing gear design. The text is backed up by calculations, specifications, references, working examples, and nearly illustrations. Advanced Search. Year Principles and Practices. Procedures for other aircraft types are given later, but the process is similar in all cases.
The next step is to select tire sizes, but this cannot be done until the static loads have been determined. It is usual, however, to locate the nose landing gear at this stage.
Conversely, the load should not be too light; in that event, steering would be difficult and the righting moment in a drift landing would be marginal. From a review of the structure, a suitable support frame must be deter- mined, preferably so that the gear will retract forward, as illustrated in Fig. The latter feature is most desirable Ifl,. The pilot merely pulls an emergency release lever that releases the uplocks and the trapped actuator fluid if used , after which gravity and air drag pull the gear into a down-and-locked position.
This capability should also be used on the main gear if possible. Having selected the appropriate support frame, the next step is to suspend the gear from it and to assume initially that the wheel center will be about 3 in.
Then, the nose gear load must be calculated. The calculation of nose gear load uses the diagram shown in Fig. When the tires are selected, at a later step, it is necessary to know the nose gear dynamic load. If the minimum static nose gear load is too small, i.
Note that very small main gear movements usually have a pronounced effect on nose gear loads. Two struts are normal CFour struts. In many cases, it is necessary to move both the nose and main gears somewhat to obtain a satisfactory overall compro- mise in the loading. It may also be necessary to deviate slightly from the deg angles used in step 6.
The nose and main gears have now been located in the side view and the static loads are known. A preliminary tire selection can now be made.
It is first necessary to decide how many tires will be used on each strut. In many cases, the answer is obvious. Table 3. All aircraft weighing ,, lb have two main gear struts and four tires per strut. Below 60, lb, it is possible to use either one or two tires per strut. If it is practical, two tires per strut should be usedwit is safer!
Between , and , lb, a decision must be made as to whether there will be two or four tires per strut. The answer is controlled to some extent by the anticipated stowage concept. For instance, the C uses two very large tires on each side of the aircraft; they are placed in tandem and the fuselage pod can be relatively slim.
If a four-wheel bogie had been used, the pod would have been f a t t e r - e v e n though the tire sizes might have been smaller. As aircraft approach , lb, runway loading becomes more impor- tant, a factor that cannot always be sufficiently alleviated by merely increas- ing the tire size or number of tires per strut. In that event, the only solution is to increase the number of struts. The Boeing and Lockheed C-5 are typical examples. Tire Selection From the maximum main gear static load previously calculated, it is nec- essary to divide that load by the number of tires per strut to obtain the static single wheel load.
Two problems have to be considered for the nose gear: These loads previously calculated are divided by the number of nose gear wheels to obtain the single-wheel static and braking loads.
With these data, it is then possible to use the tire manu- facturers' catalogs to select the tires. Typical data for tires are given in Table 3. As an example, consider an aircraft with the following characteristics: Speed Max Max Ply rating, press. That rating is chosen because it is more severe than the static rating, a feature that is discussed further in Chapter 6. The specifications require two tires on each main gear and two tires on the nose gear.
Thus, the tire loads are as follows: With this factor, the loads are as follows: The tires listed in Table 3. It is clear that several tires are capable of meeting the required load conditions. The selec- tion, then, must be based upon factors other than load. In this case, a 25 x 6. If the aircraft is a corporate jet, a 29 x 7.
Load and pressure vary almost linearly at normally considered tire deflections; thus, if psi is required for 13, lb, only psi will be required for an actual load of 10, lb. Commercial operators prefer the lower pressures in order to maximize tire life and minimize runway stresses.
The nose gear tire selected for the corporate jet would be the 26 x 6. Some of the tradeoffs involved in tire selection are discussed in later chapters.
For instance, the nose gear tire weighs With two tires per aircraft, a weight penalty of 27 lb is thus paid to obtain the lower tire pressure.
To place the tires in the deflected vertical position, note their loaded radii on the tire selection charts. For the 29 x 7. This is the distance from the ground to the axle center with the aircraft static and the tire at optimum deflection.
The nose gear tire is a bit more complicated: This allows the nose gear axle center to be determined and, as with the main gear, it becomes the starting point for determining compressed and extended shock strut positions. At this point, no further work is usually done on the landing gear in the conceptual design phase.
The tires are shown on the three-view drawing with no visible means of connection to the airframe. The static ground lines and tail-down lines are also shown, as depicted in Fig. Assume that the gear is a normal design in which the wheel and strut travel are the same. The first step is to determine the maximum load accept- able in the shock strut. This load comprises the static load plus the dynamic reaction load. When that load is divided by the static load, the reaction factor N is obtained.
This is sometimes called the landing gear load factor or merely the landing load factor. Its value ranges from 0.
Its permissible magnitude is determined by the airframe designers and struc- tures specialists. They must design the airframe to accomodate those factors during landing. Initially, the aircraft is assumed to be a rigid body, with no relative accel- eration between the c. Thus, the load factor at the c. When the aircraft comes to rest on the ground, the lift is zero and the shock strut force is equal to the aircraft weight: Therefore, Fs W No.
The methodology is as follows. Then, 2[ 0. By adding 1 in. For an initial layout, assume that a quarter to a third of the total stroke is used in moving from static to compressed. Thus, for a 9. The ground lines with gears compressed and the tail-down line and angle can now be added to the side view. The next step is to develop the basic kinematics concepts from which the "stick diagrams" are prepared. Some typical examples are shown in Fig. The possibilities are limitless, depending on the ingenuity, imagination, and know-how of the designer.
Lateral Location of Main Landing Gear The lateral location of the main landing gear affects the turnover angle and the ground clearances with movable surfaces such as ailerons and flaps, wing tips, engine nacelles, and, if used, propellers.
The dia- gram shows a twin-wheel nose gear which is different from that shown in various requirements documents where a single wheel is shown. With the latter, X" and C arc obviously zero. When there arc more than one wheel at either the nose gear or main gear, assume that the aircraft will tip along a line drawn through the outboard wheels. The angle 0 must not be more than 63 deg for land-based aircraft or 54 dog for carrier-based aircraft.
Although some aircraft do, in fact, approach these values, it is desirable to make it as small as possible. Note that it is sometimes extremely difficult to have low angles on high-wing aircraft because their landing gears arc often mounted on the fuselage side and thereby have narrow tracks.
Since short takeoff and landing STOL aircraft arc usually high-wing configurations, a high turnover angle is one of the problems the designer must solve. Y Fig. De Havilland Canada aircraft such as the DHC-5, Dash-7, and Dash-8 have nacelle- mounted gears with a consequent reduction in the turnover angle. Another approach is to use a bicycle-type gear, as on the B, with outrigger wheels between the siamese engine nacelles to restrict turnover.
This is illustrated in Fig. I0 depicts a method that was used by the author some years ago on a design that did not proceed beyond the study stage. The gear is sus- pended from the rear spar of a high wing and retracts forward into a stream- lined pod. The above values were calculated by the author and may vary somewhat from manu- facturers' calculations due to differences in assumed critical center-of-gravity positions.
Clearance Checks There is now sufficient information to enable clearance checks to be made. This is where present-day computer graphics arc particularly valuable.
In a nutshell, these checks involve placing the aircraft in all the worst attitudes possible, with several landing gear failure conditions, and then checking to see if there are still adequate clearances with all moving and fixed parts of the aircraft. The results of these analyses often require changes to Ix: These changes can include rcfairing the aft fuselage, moving or shortening belly antennas, moving the engines inboard or upward, restricting control surface deflection, and lengthening the landing gear or moving it outboard.
IF- m z I Are the nose and main gear shock struts operating properly and the tires at normal inflation? Is the main gear shock strut fully compressed and the tire flat on one side, with static deflection and normal tire inflation on the other side? Is the main gear shock strut on the other side fully compressed and the tire flat? They are summarized in Fig. No attempt is made to define detail requirements on parts that are normally provided by vendors, e.
Also, the source of a requirement is not given whenever it is considered to be acceptable internationally and by both military and civil authorities. In a few cases, U. Navy requirements are peculiar and these are noted; also British requirements are slightly different in some areas and these too are highlighted.
This allows follow-on sales of a commerical vehicle, for instance, or a derivative of it to military customers or to foreign countries. The penalties paid are often minor if these requirements are considered initially.
As an example, some agencies require the main landing gears to be interchangeable left and fight. This is obviously a benefit, so the feature should be incorporated whenever possible, whether it is required or not. The specifications cited in this chapter are listed in Chapter AIR provides a complete page dic- tionary-like listing of terms that are used in landing gear designDtend- ing to reinforce those critics who proclaim that landing gear designers have their own language!
It is, however, an extremely useful compendium of terminology that should be studied by anyone who is seriously involved with this subject. Unless there are other means to decelerate the aircraft in flight at speeds up to 1. It should be possible to retract and extend the landing gear satisfactorily under the most adverse flight conditions occurring throughout the range of airspeeds from Vso to VLO and accelerations of 0. A list of typical airspeed limits is provided in Table 4. Air Force and with Fig.
In both cases, at the design gross weight, the designer must ensure that the plane of each wheel is vertical. Much of this information is based on specifications developed cooperatively by industry, government, and various engineering societies see Chapters I and The word "shall" is used in such specifications to denote a definite requirement and is thus repeated here.
Aircraft Landing Gear Design: Principles And Practices
Navy landing gear layout requirements source: Navy Specifica- tion SDJ. The report that accompanies these drawings shall indicate tire sizes, tire inflation pressures, design sink speeds, total air volume, and isothermal pressure in the extended, static, and compressed shock absorber positions at maximum takeoff gross weight, as well as the preliminary loads imposed upon the landing gear USN. The strut shall be designed to use MIL-H hydraulic fluid. The air connections shall conform to AND and the air valve shall be in accordance with MS All struts shall incorporate MS scraper tings installed per MS The packing gland nuts at the end of the shock absorber shall have wrench slots as defined in MIL-S Minimum chrome plating thick- ness shall be 0.
To demonstrate that there is sufficient oil above the orifice to avoid foaming, two successive drops shall be made within 5 min and then repeated after removing oil corresponding to 0. At the threads for the wheel bearing retainer nut, there shall be two cotter pin holes at 90 deg spacing.
For normal, utility, and aerobatic category aircraft being certificated by FAR Part 23, the sink speeds and wing lift are determined by formulas given in Part The wing lift, for instance, cannot be more than two-thirds of the aircraft weight at touchdown, the inertia load factor cannot be less than 2. For all other types, refer to FAR Part 25, paragraphs Some of the requirements are summa- rized below: If these measurements are made by drop tests, the free drop heights must not be less than Refer to FAR At the design landing weight, the sink rate shall be 5.
Demonstrate that there is sufficient capacity to withstand landing at 1. The tire and wheel flange dimensions are taken from the manufacturers' catalogs. Usually, the tires are inflated to pressures less than the maximum rated values listed in the catalogs and MIL-T; in that event, the pressures are reduced linearly with load.
Where twin tires are used, inflate to equal pressures. This load factored by 1. This load, with 1. Nose wheel tire inflation pressures are based on maximum allowable dynamic loads.
This may be accomplished, for instance, by adding plies to the tires. On a multiwheel main gear, design it so that if one tire or wheel fails during taxi or takeoff at the maximum gross weight, the remaining tires and wheels on that gear can withstand the most severe overload conditions imposed. Determination of this overload must be based on an elastic analysis of the aircraft and all parts of the landing gear.
Navy aircraft, the ply rating shall be at least two plies less than the maximum rating recommended by the Tire and Rim Association. On carrier-based aircraft, the operating pressure shall not exceed 1. A minimum tire section width of 6 in. In addition, unless the tire is prevented from spinning during retraction, an extra 2. Wheels are designed in accordance with MIL-W The aircraft manufacturer is responsible for calculating the maximum static and dy- namic loads on the wheels, which must be less than their rated loads.
MIL-W also calls for the installation of thermal-sensitive pressure- release devices fuse plugs , such as depicted in Fig.
A means must be provided to prevent water from entering the wheel bearings. Static test the combined wheel and tire to a pressure equal to 3. Consider the use of nonfrangible wheels to prevent airframe damage due to wheel disintegration after tire failure. This wheel type is capable of rolling for a specified distance without shedding any pieces capable of piercing the airframe.
Clearances were shown in Fig. Navy carrier-based aircraft, the centers of the main wheel axles must clear the deck by at least 6.
USAF requirements stipulate, and BCAR recommends, that wheels shall be stopped from rotating during retraction or prevented from rotating in the retracted position. This prevents parts inside the wheel well from being damaged by the flailing of a damaged tire and prevents undesirable wheel-rotation noise from being transmitted to the crew and passengers.
It may be accomplished by braking the wheels or by friction pads. FAR Part 25 does not require this, but it does require that a loose tire tread must not cause any damagemwhich may amount to the same thing!
Brake control systems are designed in accordance with MIL-B The U. Navy requires the use of method II SD, paragraph 3. This is the greater of 1. Usage of these two capacities is reflected in the tests conductedm95 stops at normal capacity and 1 stop under the emergency condition.
Wheel brake capacity requirements for U. For military aircraft, there are two published tables of wheel brake capacity require- ments: The former, shown in Table 4. In that event, the requirements of AIR may be more appropriate since they were determined by SAE experts representing both industry and government and reflect current thinking. See Table 4. The wheel brake field service life spectrum illustrated in Table 4.
There are many details to be recognized, but to provide the reader with a suitable example, the following is an abbreviation of the two methods for evaluating general performance and wear: At least 25 stops should be made with kinetic energy equal to the certified normal brake energy capacity, with constant brake pressure at the normal value.
Measure the stopping time for each fUll. At least 5 stops should be made as shown in method 1, plus 95 similar stops which may be made at reduced speed to allow maximum brake usage in stopping the aircraft from the greater of the following: In addition to the above, the BCAR includes tests for static force, reduced speed stopping, overload, and certified emergency brake energy. Flexible lines should be routed so that brake heat cannot cause them to rupture.
Locate all brake lines on the aft side of the shock strut so that they are protected from foreign object damage. Provide an emergency system capable of stopping the aircraft in the same distance as the normal system.
The emergency system shall be completely independent of the normal system upstream of the brake shuttle valve or its equivalent. If drag chutes are used to augment deceleration, they should be in accordance with MIL-D In addition to the above, the FAR has the following requirements.
AS For other than friction materials, the assembly shall withstand the 65 stops without failure or impairment of operation. Rotorcraft speed at brake application shall be determined by analysis. The assembly shall withstand the test without failure and without impairment of operation, for other than friction materials. The average deceleration shall not be less than the average noted in Table 4.
The aircraft must have a p a r k i n g b r a k e that, w h e n set by the pilot, will pevent the aircraft from rolling on a paved, level r u n w a y with t a k e o f f p o w e r on the critical engine. In addition, B C A R requires that b r a k e forces m u s t increase or decrease progressively as the force or m o v e m e n t is increased or decreased at the brake control. ARP CThe dynamic torque sequence shall be conducted with 3 sequences of 10 land plane landing design gross weight stops followed by 1 maximum landing gross weight stop.
CA new brake shall be used for the rejected takeoff RTO stop. This brake may be conditioned prior to the RTO demonstration. General Notes: I The calculations for capacity requirements shall represent the worst situation that affects overall sizing of the brake. CThe worn brake RTO stop is conducted to determine the aborted mission KE capacity of a worn brake and to demonstrate the ability of the brake to complete an aborted mission stop to reasonable conditions.
See general note 1. General Notes 1 The analysis is to be based on realistic average conditions expected to be experienced in service usage of the aircraft. The brake will be refurbished with a new complement of disks or other heat sink members, linings, and seals.
Cooling air of 30 knots may be used during all taxis. Taxi snubs during rolling may be specified if applicable to the aircraft system. If arresting hooks are used for deceleration, requirements pertaining to their installation are shown in MIL-A- FAR Part 25 and the USAF requires that the system must be designed so that no single failure will result in a hazardous loss of braking capability or directional control of the aircraft.
The USAF requires adequate ground control when landing on wet or icy runways or with strong crosswinds. Also, all aircraft that touch down above knots must be equipped with antiskid brake control systems, although deviations will be granted if the contractor can prove that they are unnecessary.
The BCAR requires that antiskid devices be no less reliable than the rest of the braking system, that a warning be provided to the crew to show failure of the electrical power supply to the system, and that, if any part of the system malfunctions, the affected brake units will automatically revert to a control ensuring no hazardous loss of braking or directional control. Navy aircraft have an additional requirement SD that the nose wheel shall swivel through deg without manually disconnect- ing the steering linkage.
The BCAR stipulates that, after extension of the gear and prior to touchdown, the nose wheel shall be automatically positioned in a fore-and-aft attitude; or, if it is otherwise positioned, it will neither be overstressed nor cause any hazardous maneuver.
No exceptional skill must be required to steer the aircraft, including the conditions in crosswind or sudden power unit failure. Design the nose gear towing attachments so that no damage will be caused on the nose wheel assembly or steering assembly.
In a powered steering system, the normal power supply for steering shall continue without interruption if any one power unit fails.
At ground idling, the remaining power unit s shall be capable of completing an accelerate-stop and a landing rollout. In addition, no single fault shall result in a hazardous maneuver.
The system must have sufficient torque to turn the steered wheels through their full steering angle without requiring forward motion of the aircraft or asymmetric engine thrust. This capability must be available throughout the design temperature range, at critical weight and c.
These ranges shall not require any manual disconnects unless authorized by the customer; automatic disconnects are allowed, provided that they re-engage automatically when the wheel re-enters the power steering range. The system shall provide dynamic and damping stability for all ground speeds up to 1. The shimmy requirements shall be determined by a nonlinear dynamic analysis that recognized deadband, friction, wheel un- balance, and damping characteristics.
The system shall provide sufficient damping to reduce shimmy oscillation amplitude to one-fourth or less of the original disturbance after three cycles. The BCAR requires that the nose wheel should be capable of free castoring while on the ground.
Also, the engagement of any locking devices should not restrict that capability. This document also specifies that, unless the nose wheel is automatically centered when lowered, tests must be made to prove its satisfactory functioning when the aircraft is landing with the nose wheel offset at its maximum possible angle. Class A: Class B: They must be capable of preventing retraction or extension under all loads applied to the gear.
It is not permissible to hold the gear in the up position by using door locks; the gear must not rest against the doors at any time.
An uplock must not be dependent upon proper servicing of the shock absorber. Hydraulically operated locks must not be capable of unlocking due to pressure variations and electrically operated locks must not unlock due to any faults in the electrical system.
Downlocks are generally not allowed to be stressed by ground loads, but when this is unavoidable they must have adequate strength, be nonadjustable, and be easily inspectable. A ground safety lock is required on each retractable gear, which should be lightweight, quickly releasable, installed manually, easily removable, and incapable of being installed incorrectly. Navy aircraft, it is further required that whenever overcenter links are used, a positive integral mechanical lock shall be provided at the knee.
Down-and-locked position switches shall be actuated directly by the lock. Rigging of locks shall be simple and devoid of close-tolerance adjustments. Commercial aircraft requirements state that there must be a positive means to keep the gear extended, in flight and on the ground. However, it is normal practice to apply most of the military requirements to commercial aircraft.
Table 4. The landing gear should be operable for at least cycles using the normal system and cycles using the emergency system. Do not use cables or pulleys except in emergency systems. A gravity system is preferable, assisted if necessary by a spring. Do not use a system that requires hand-pumping by the pilot.
Do not use telescoping rods or slotted links. If these systems are electrically operated, use rugged switches that will not ice-up and mount them on rigid supports to prevent malfunctions due to bracket deflections or the presence of foreign matter. Also, ensure that the gear can be extended if an electrical circuit fails. Navy aircraft, the gear shall be retractable in not more than l0 s. A safety lock is required to prevent retraction when the aircraft is on the ground and an over-ride must be provided to enable the pilot to bypass this lock if conditions warrant it.
If a touchdown switch is used to provide this safety lock, then it must operate when the main gear has compressed not more than l in.
Navy aircraft are required to be able to extend the gear in 15 s or less and an emergency system must be provided to extend the gear if any part of the normal system, or its power supply, fails. A gravity system is preferred for this purpose, with direct mechanical release of the locks.
FAR Part 25 has detailed requirements on aural warning devices and on switches to actuate position indicators. A red light is illuminated whenever the gear is not down and locked and when the gear, its doors, and its selector are in the retracted position.
Cockpit controls for steering systems are provided in MIL-S The brake control system specification MIL-B requires a warning light to show any brake system malfunction; the parking brakes were discussed previously in Sec. Controls are also required to engage or disengage the antiskid system if used and also to set the degree of braking if an automatic brake system is employed. Navy requirements SD note that emergency landing gear control shall be separate from, but as close as practical to, the normal control unless approved otherwise.
The design must preclude interaction between normal and emergency operation; the failure of the normal control must not impair actuation of the emergency system. MIL-L includes the following suggested requirements that are associated with protection: Navy requirements. ASFC DH requires shock struts, forks, and axles to be designed so that mud will be prevented from entering internal cavities. Special care should be taken to plug the axles so that mud cannot contaminate the bearings.
Navy also requires that the fairings design shall preclude the accumulation of mud, dirt, or cinders. Exposed mechanisms, equip- ment, electrical wires, and fluid lines should be positioned so that they will not be damaged by foreign objects thrown from the tires.
It is suggested that one partial solution is to close the landing gear doors after gear extension and to provide easily removable covers to exposed parts. The U,S. Navy requires that any wheels and tires that are retracted into a position close to a heat source must be protected from that heat.
BCAR requires that brakes be protected from the ingress of any foreign matter that may impair their proper functioning. It has a similar require- ment to those stated above concerning the effects of burst tires and wheels.
Navy aircraft, doors that close after gear extension should be designed so that they can be opened from the ground. Further guidance is provided in AFSC DH, which advocates that all hydraulic mechanisms have their filler plugs, bleeder plugs, and air valves placed for easy servicing.
Design shock struts so that it is possible to determine the extent of its inflation by using only a scale. Prepare the interior of the wheel wells with a MIL-P primer coating. It should be possible to remove a wheel without removing any other part of the gear and the jack pads should be so located that the jacks will not affect operation of the gear.
Details of these conditions are too voluminous to be included here and reference should be made to the specification. AFSC DH notes that the design of a multiple-wheel gear should be such that, if one tire or wheel fails during a maximum weight takeoff, the remaining tires and wheels on that gear can absorb the severest overload conditions imposed.
This overload is determined by an elastic analysis of the aircraft and its landing gear. This specification defines the location of the hook, the obstacles to be overcome on the carder deck, the design of the hook itself, its installation details, the applied loads, the controls associated with the hook, and the requirements pertaining to its shock absorber.
The Appendix to this specification shows how to determine the aircraft pitch attitude. Some do not have tires, wheels, brakes, antiskid devices, retraction systems, or steering systems, but all of them have some form of shock absorber. While the carrier landing has sometimes been called a "controlled crash," it would be a complete catastrophe without the shock absorber. Since this part is undoubtedly the most important component in the landing gear, this chapter will discuss it in considerable detail.
The basic function of the shock absorber, or shock strut as it is often called, is to absorb the kinetic energy during landing and taxiing to the extent that accelerations imposed upon the airframe are reduced to a tolerable level. The gas is usually dry air or nitrogen. Figure 5. In selecting the type, due recognition must be given to the simplicity, reliability, maintainability, and relatively low cost of the solid-spring shock absorbers.
On smaller utility aircraft, the weight penalty is usually negli- gible and the noted advantages far outweigh the penalties in such cases. A simplified procedure is included later in this chapter to calculate the characteristics of this type of gear.
Rubber Springs Shock absorber efficiency is dependent upon the degree to which the shock-absorbing medium is uniformly stressed. These are vulcanized to plates and are stacked as shown in Fig.
In order to permit satisfactory vulcanizing, each disk is generally no more than 1. They have been widely used--the Twin Otter design shown previ- ously is an example. Further details of designing with rubber blocks are given later in this chapter. British Aerospace. Dowty Rotol Ltd. They are similar in design to the oleo-pneumatic shock absorber, but are heavier, less efficient, and less reliable and have no inherent means of lubricating the bearings.
Since they are not used today, no further details are given here. Oil The so-called "liquid spring" Fig.
They are still used today, mostly in levered-suspension designs. Its advantages are low fatigue due to the robust construction and relatively small size. Its disadvantages are the fact that fluid volume changes at low temperatures affect shock absorber perfor- mance, the shock absorber can be pressurized only while the aircraft is on jacks i. Typical calculations are provided later in this chapter.
Internally Sprung Wheels Although these are no longer in use, the concept is interesting enough to warrant documentation in this section. The internally sprung wheel was developed by Dowry in the 's and was used on the Gloster Gladiator.
It is shown in Fig. Its advantages were that it enabled a rigid leg to be used, but its disadvantages were that a large tire was needed to match the large wheel required for reasonable shock absorber travel; also, the diffi- culty in accomodating a brake is obvious. In addition, the available stroke is really too small for contemporary aircraft.
They have the highest efficiencies of all shock absorber types and also have the best energy dissipation; i. In the design shown in Fig. I i,--, When the aircraft lands, oil is forced from the lower chamber to the upper chamber through the orifice.
Although this need only be a hole in the orifice plate, the hole area is often controlled by a varying-diameter metering pin, as depicted in Fig. Typical calculations for an oleo-pneumatic strut are provided later in this chapter. Some finer points will emerge as the design progresses--such as whether the design will specifically prevent mixing of the gas and oil. In practice, sink speeds of this magnitude are very rarely achieved. These types of aircraft normally approach at a deg glide slope.
Short takeoff and landing STOL aircraft are designed to approach at a higher angle deg and to minimize flare. Load Factor Load factors applied to the landing gear should not be confused with aircraft load factors. The latter result from maneuvers or atmospheric disturbance. The landing gear load factor is, to some extent, a matter of choice, the details of which are given in Chapter 3. From this, the approx- imate available load factor can be obtained and used in the overall structural analysis.
From this and subsequent iterations, the landing gear load factors are prescribed by the structures department. In many cases, the airframe design will not be controlled by the landing load factor, except for localized areas adjacent to the gear.
The author was involved in such a design a STOL aircraft where the aft fuselage loads were controlled by the high empennage loads and most of the wing was controlled by gust, flap, and aileron loads.
Aircraft Landing Gear Design: Principles And Practices
Only the wing engine mounts were affected by the landing gear loads. This does not equal geometric area. A perfectly conducting sphere of projected cross sectional area 1 m2 i.
Note that for radar wavelengths much less than the diameter of the sphere, RCS is independent of frequency. Modern stealth aircraft are said to have an RCS comparable with small birds or large insects,  though this varies widely depending on aircraft and radar.
If the RCS was directly related to the target's cross-sectional area, the only way to reduce it would be to make the physical profile smaller. Rather, by reflecting much of the radiation away or by absorbing it, the target achieves a smaller radar cross section.
Enemy radar will cover the airspace around these sites with overlapping coverage, making undetected entry by conventional aircraft nearly impossible. Stealthy aircraft can also be detected, but only at short ranges around the radars; for a stealthy aircraft there are substantial gaps in the radar coverage.
Thus a stealthy aircraft flying an appropriate route can remain undetected by radar. Even if a stealth aircraft is detected, fire-control radars operating in C, X and Ku bands cannot paint for missile guidance low observable LO jets except at very close ranges.In support of this definition, landing gear design is considered to include the following items: Initial Layout 3.
These options will be reviewed to see how they match the airframe struc- ture and the flotation requirements if any. Those cited here are also included in Chapter In this analysis, the basic energy equation still applies, although the tire effect can probably be ignored for simplicity.
As aircraft size increases, it is often necessary to have a variable orifice.
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