Construction of a retractable single-strut undercarriage in fiber composite design for the motorized sailplane “Valentin Taifun 17E”

Content

I. LIST OF TABLES

II. LIST OF FIGURES

III. LIST OF SYMBOLS

IV. ABSTRACT

1. INTRODUCTION AND PRESENTATION OF THE TYPE OF PROBLEM
1.1 Valentin Taifun 17E
1.2 Standard carriage of the Taifun 17E, its problems and task of the thesis

2. FUNDAMENTALS OF THE MANUFACTURE OF SANDWICH-STRUCTURED COMPOSITES
2.1 Fundamentals
2.2 Showcase of manufacture of a sandwich constructed fuselage plate

3. BOUNDARY CONDITIONS
3.1 Determination of position of the landing gear in the extended position
3.1.1 Determination of center of mass under consideration of different load conditions
3.1.2 Determination of the horizontal landing gear position under disregard of plane inclination
3.1.3 Comparison with available space
3.1.4 Adaption of the landing gear position
3.1.5 Provision for the start-kick
3.1.5 Further considerations

4. LOAD ASSUMPTION OUT OF THE JOIN AVIATION REQUIREMENTS 22
4.1 Level landing condition
4.2 Lateral load conditions
4.3 Lever control
4.4 Limit maneuvering load factor
4.5 Wheel
4.6 Factors of Safety

5. CONSTRUCTION POSSIBILITIES
5.1 Version 1 - Description and Evaluation
5.2 Version 2 - Description and Evaluation
5.3 Version 3 - Description and Evaluation

6. CONSTRUCTION AND STRESS ANALYSIS VERSION 3
6.1 Wheel
6.1.1 Experimental set-up
6.1.2 Calibration
6.1.3 Execution
6.1.4 Selection of the tire pressure
6.2 Axle
6.2.1 Forces on the axle by “normal landing load case”
6.2.2 Forces on the axle by “lateral load case”
6.2.3 Contact pressure between axle and fit bushing
6.3 Bushings
6.3.1 Execution of the bushings
6.3.2 Bearing Pressure between spherical plain bearing and bushing
6.3.3 Bearing pressure between bushing and roving truss
6.3.4 Implementation of the brake
6.4 Gear swing
6.4.1 Framework
6.4.2 Bending truss
6.4.3 Bending truss with shear field
6.3.4 Compression strut - behavior during compression
6.5 Framework trusses 1 and 2
6.5.1 Framework truss 1
6.5.2 Framework truss 2
6.5.3 Console wall
6.6 Connection of the framework trusses 1 and 2
6.7 Connection of framework truss 1 with the wheel and the gear swing
6.8 Bolt mechanism
6.8.1 Requirements of the single components
6.8.2 Execution and assembly
6.8.3 Improvements
6.9 Lever mechanism - extending and retracting
6.9.1 Requirements and execution
6.9.2 Actuating force
6.10 Bench face - Holm mounting
6.10.1 Contemplation and determination of loads transferred by axle
6.10.3 Thickness for perpendicular/lateral loads
6.10.4 Adhesive bonded joint
6.11 Connection of the holm mounting with the gear swing and framework truss 2
6.11.1 Manufacturing and assembling
6.11.2 Stress analysis of the “bench face axle”

7. EVALUATION AND FURTHER PROCEDURE

LIST OF REFERENCES

V. APPENDIX

V.1 Overdetermined system

V.2 Spring characteristics - Results

V.3 Component drawings
V.3.1 Landing Gear
V.2.2 Bolt Mechanism
V.2.3 Fuselage and Bench Face

I. List of tables

Table 1: Listing of part masses and their lever arms

Table 2: Set-Up for Xf-max

Table 3: Set-Up for Xf-min

Table 4: Resultant horizontal position Lh for the landing gear

Table 5: Calibration table

Table 6: Spring constants for different pressures

Table 7: Deflection for different pressures

II. List of figures

Figure 1: Valentin Taifun 17E

Figure 2: Coating the counterpart with resin

Figure 3: Foam plate on the glass fiber layers

Figure 4: Connection of the hose with the “film chamber”

Figure 5: Finished sandwich element

Figure 6: Reference line

Figure 7: Determination of balance point

Figure 8: Reference line HE

Figure 9: Position of the wheel

Figure 10: Position of landing gear mounting

Figure 11: Final position of the wheel

Figure 12 : Geometry of plane on ground

Figure 13: Start kick

Figure 14: Maximal Tail-Skid Length

Figure 15: V1 Retracted position

Figure 16: V1 Over-twisting of the toggle struts

Figure 17: V1 Extended position

Figure 18: V2 Retracted position

Figure 19: V2 extended position

Figure 20: V3 retracted position

Figure 21: extended position

Figure 22: Experimental set-up

Figure 23: Maximal deflection of the tire with 0.5 bar

Figure 24: forces on the axle - side view

Figure 25: loads on the axle by "normal landing load" - front view

Figure 26: overdetermined axle system

Figure 27: "TM-Interaktiv" - Spherical Plain Bearings

Figure 28: Central system of forces - determination of force Fft

Figure 29: Axle fixed by plain bearing bushes

Figure 30: TM-Interaktiv "Normal Load Plane Bushing"

Figure 31: Reaction forces by lateral load

Figure 32: "TM-Interaktiv" - Lateral load case

Figure 33: Contact pressure cylinder - cylinder

Figure 34: Contact pressure by moment

Figure 35: Intersection B of bushing

Figure 36: Intersection D of bushing

Figure 37: Bushing with break attachment

Figure 38: Framework gear swing

Figure 39: Deflection bending truss

Figure 40: Shear field system

Figure 41: Landing gear shear field - sectional view

Figure 42: Landing gear shear field - front view

Figure 43: Shear buckling

Figure 44: Framework Truss 1 - Load cases

Figure 45: Manufacturing of framework truss

Figure 46: indeterminate straight bumper - bolt

Figure 47: Dotted function lines and direction of the forces - bolt

Figure 48: Shear force due to strap collocation

Figure 49: Guide bushes with flange

Figure 50: Strap Collocation - Transfer of loads into a plane structure

Figure 51: Position of the steering trusses

Figure 52: Position of the screws in the bushing of framework truss

Figure 53: Virtual position of the screws in framework truss

Figure 54: Gear-axle mounting

Figure 55: Extented bolt position

Figure 56: Retracted bolt position

Figure 57: Framework truss without plate

Figure 58: Framework truss with plate

Figure 59: Wedge on the truss at the bushing

Figure 60: Thrust lever with guide pulleys

Figure 61: Mounting "Framework truss 2 - Linking Truss - Pneumatic Spring"

Figure 62: Lever mechanism - extended position

Figure 63: Lever mechanism - retracted position

Figure 64: Determined abutting face of the landing gear

Figure 65: “Bansbach easylift Typ B” with Bowden cable

Figure 66: System of forces for lever mechanism

Figure 67: Forces in bench face caused by Fft

Figure 68: Forces in bench face cause by FB,n

Figure 69: Central force system - bench face axle

Figure 70: Enhancement of the bench face with ribs

Figure 71: Connection of the axle with the bench face

Figure 72 : Complete Landing Gear with Lever Mechanism

III. List of symbols

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IV. Abstract

The purpose of this thesis is to calculate, interpret and construe a retractable one-wheel landing gear for the motorized sailplane “Valentin Taifun 17E”. Therefore the available space in the sail plane was measured and the data was transferred into the CAD-system “Autodesk Inventor”. The calculative position of the extended landing gear, in dependence of the variable center of mass and in due consideration of the requirements of the JAR 22, was determined. Due to the information about the position of center of mass and of further considerations, the position of the landing gear mounting could be determined so the retraction kinematic of the carriage was automatically stated.

The extension should be realized with lever mechanism and the carriage needs to be locked in extended position. With the aid of a market research and creativity techniques, different construction possibilities were created. These constructions were assessed, so that a winner version could be chosen for further contemplation.

In order to construe the single components of the carriage it was necessary to identify the occurring forces and to design and construe the components in dependence of the characteristic cases of damage. To determine the acting forces it was necessary to take into account the spring constant and the optimal tire pressure of the used tire, which was therefore determined in the course of this diploma thesis. The construction process was done according to the flow of forces during the landing impact. During the construction many different components were developed. They were compared with each other, so that the best part accomplishes the most with the requirements: lightness, stability and simplicity.

The developed construction needs to be considered as a first prototype and as the basic for further improvements of the carriage, before it is getting finally implemented into the “Valentin Taifun 17E”.

Kurzfassung

Diese Diplomarbeit hat das Ziel, ein einfahrbares Einbeinfahrwerk für den Reisemotorsegler „Valentin Taifun 17E“ auszulegen und zu konstruieren. Dazu wurde im Vorfeld der vorhandene Flugzeugrumpf ausgemessen und in das CAD-Programm „Autodesk Inventor“ eingepflegt Unter Berücksichtigung der Bauvorschrift „JAR 22“ wurde die Position des ausgefahrenen Fahrwerks, in Abhängigkeit des Schwerpunkts, bestimmt. Im Anschluss konnte die Position der Fahrwerksbefestigung im Rumpf bestimmt werden, sodass die Ein- und Ausfahrkinematik der Fahrwerksschwinge festgelegt war.

Der Ausfahrvorgang soll mit Hilfe eines Hebels realisiert werden, sowie das Fahrwerk im ausgefahrenen Zustand fixieren. Mit Hilfe einer Marktrecherche und unterschiedlichen Kreativitätstechniken wurden mehrere Konstruktionen erstellt. Diese wurden miteinander verglichen, sodass eine Siegervariante für die weitere Betrachtung ausgewählt werden konnte.

Um die einzelnen Bauteile des Fahrwerks, unter Berücksichtigung der wirkenden Belastungen auszulegen, mussten zuerst die wirkenden Kräfte bestimmt werden. Dazu wurde in einem Versuch die Federkonstante des vorhandenen Reifens für verschiedene Drücke ermittelt und daraus resultierend der optimale Reifendruck bestimmt. Die Auslegung der einzelnen Komponenten erfolgte in der Reihenfolge des Kraftflusses. Es wurden mehrere Konstruktionsvarianten für die einzelnen Bauteile und Bauteilgruppen erstellt und miteinander verglichen. Dabei wurde darauf geachtet, dass die Anforderungen: geringes Gewicht, Festigkeit und Einfachheit bestmöglich erfüllt werden.

Das so erstellte Fahrwerk dient als erster Prototyp, auf dessen Basis weitere Verbesserungen durchgeführt werden sollten, bevor es in den Reisemotorsegler „Valentin Taifun 17E“ eingebaut wird.

1. Introduction and presentation of the type of problem

1.1 Valentin Taifun 17E

The “Valentin Taifun 17E” is a composite designed motor glider with two juxtaposes seats and a T-empennage, which was built from 1981 to 1990 by the company “Valentin GmbH” in Königsbrunn. In contrast to the competing companies, “Valentin GmbH” decided to equip the Taifun 17 E with detachable wings, which can be folded along the fuselage for better storage and transportation and a retractable tricycle landing gear for among other things a better aerodynamic performance. The aircraft has a wingspan of 17 meter and is equipped with a “Limbach L2000EB” motor with 60 kW (80 hp)^[1].

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Figure 1: Valentin Taifun 17E ^[2]

1.2 Standard carriage of the Taifun 17E, its problems and task of the thesis

The Valentin Taifun 17E was usually equipped with a tricycle landing gear. It consists of two main gear wheels and one nosewheel. Optional it was possible to equip it with a tail skid wheel. The nosewheel can be used for steering and it has a special kinematic retracting mechanism. The wheel is therefore twisted and distorted stored in the fuselage. Another special feature of the retraction system is the winder mechanism, whereby the pilot extends and retracts the landing gear. This is accomplished in the style of the winder mechanism of an old photo camera. Unfortunately this mechanism was quite complicated and therefore quit prone to failure.

Task of the diploma thesis is to construct a new landing gear as easy as possible in its kinematics, production, assembly and maintenance. A landing gear that accomplishes with these requirements the most is a one-wheel landing gear. The advantages of a one-wheel landing gear are obvious. Due to its more greatly construction it is still lighter and the handling while landing is easier. There are less used parts and therefore the retraction/extension mechanism can be easier executed. Also the production, assembly and maintenance benefit from fewer parts. Furthermore to save more weight the landing gear should be predominantly made of fiber composite materials.

2. Fundamentals of the manufacture of sandwich-structured composites

2.1 Fundamentals

A structural sandwich is a special form of composite; it consists of several materials with different properties which are compounded together under usage of a technical resin. This compound composite has a better stiffness and stability than the single components.

Typically a sandwich composite consists of four parts; two thin, force conducting top layers, which are kept away by a light, relative weak core material and a technical resin which bounds everything together.

The core can consist of foamed material, paper combs, mineral wool or balsa wood. The top layers usually consist of fiber filaments, but sometimes also steel sheets or ply wood are used.

If this sandwich material is loaded by a bending moment it can be assumed that the layers will be loaded by tension and compression forces, therefore the bending stresses in the core are neglected. The core then mainly responsible for transferring the shear loads.^[3]

2.2 Showcase of manufacture of a sandwich constructed fuselage plate

To implement the new landing gear it is necessary to cut the fuselage open. Therefore it is necessary to close the fuselage at some spots. For these closures a fuselage plate is made. In the following it is exemplary explained how to produce a plate in sandwich construction method. This method is also used at well-established factories and manufactures for sailplanes.

To produce a sandwich plate for the plane following materials are used:

- glass fiber filament
- foam plate
- laminating resin (MGS L 285)
- hardener (MGS 285)

For the production following auxiliary means are necessary:

- vacuum pump
- plastic film
- flannels
- counterpart of the fuselage
- wax (BF 700)

Step 1

The counterpart of the fuselage is waxed with BF 700, so that the finished sandwich plate can be easily released from the counterpart. For that several coatings are embrocated in an interval so that the fresh rubbed in coat can dry out before the next wax coat is applied.

Step 2

The laminating resin and its hardener are put together under a mix ratio of approximately 2.5 : 1 (laminating resin : hardener). The resin is stirred till it's totally mixed.

Step 3

The counterpart is coated with the resin and covered with the first layer of trimmed glass fiber filament. Than it is also coated with resin and a second glass fiber layer is applied under an angle offset of 90°.

Abbildung in dieser Leseprobe nicht enthalten

Figure 2: Coating the counterpart with resin

Step 4

On a desk, overdrew with a plastic film, a second layer of glass fiber is made, like as in Step 3 explained.

Step 5

The foam plate is punctured with a nail stamp, so that the resin can be easier aspired by the foam. The foam plate is laid on the layer on the counterparts with a respect of 3-4 cm to the edge.

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Figure 3: Foam plate on the glass fiber layers

Step 6

The additional layer of Step 4 is put on the foam plate. Overlaying material is cut off and flannels are put on the second layer. These flannels are responsible for aspiring excessive resin. Another plastic film is put over the sandwich element and hermetically sealed with adhesive tape, so that a “hermetic chamber” arises. A small whole is punched into the middle of the plastic film and a connecting hose for the vacuum pump is added.

Abbildung in dieser Leseprobe nicht enthalten

Figure 4: Connection of the hose with the “film chamber”

Step 7

The vacuum pump is switched on, so that a vacuum appears in the hermetically sealed chamber and the ambient pressure acts on the sandwich plate and compresses it. The resin totally infiltrates into the foam plate. The pump will be turned on for 24 hours till the resin is totally dried out.

Afterwards everything is disconnected from the sandwich element and it can be used for further converting or installation.

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Figure 5: Finished sandwich element

3. Boundary Conditions

3.1 Determination of position of the landing gear in the extended position

Before the construction process of the carriage starts, the exact position of the retracted wheel needs to be determined. The position of the retracted wheel dictates the loads at the main wheel and at the tail wheel. In reference to that, first the center of mass of the plane, in dependence of the variable masses, is located. Afterwards the vertical and horizontal positions of the extended gear are preselected with the aid of the requirements which are stated and demanded in the “JAR 22”. This preselected position is compared with the available space in the fuselage. Then the final position of the landing gear in extended and retracted position is determined and different loading cases are considered.

In the following all distances refer to the reference line which is further abbreviated with B.P. (“BezugsPunkt”). The reference line begins at the root rib of the wing (in Figure 6 it is marked with red color). For the calculative considerations the position of the root rib is determined with the aid of the original entire component drawing.

Abbildung in dieser Leseprobe nicht enthalten

Figure 6: Reference line^[4]

3.1.1 Determination of center of mass under consideration of different load conditions

The allowable center of mass during the flight, for the Taifun 17 E is between 400 mm till 540 mm, rearward the reference line (B.P.). That is about 33 to 45 % of the aerodynamic wing depth of 1.105 m.^[5] For determination of center of mass, every partial mass G has to be multiplied with its lever arm X (equation (1)). In equation (3) the lever arm of the center of mass, referring to the reference line is determined by dividing the sum of the resultant moments by the total mass (equation (2)).

The subscription L describes the characteristics of the tare weight; subscription P the characteristics of the crew; the subscription K the characteristics of the fuel and the subscription G describes the characteristics of the luggage.

The partial masses and its lever arms, stated in theJAR-22 and in the “Flughandbuch Taifun 17E”. The position of the different part masses and their positions, relating to the reference line B.P., are exemplified in Figure 7.

Abbildung in dieser Leseprobe nicht enthalten

Table 1: Listing of part masses and their lever arms^{[6] [7]}

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Figure 7: Determination of balance point^[8]

For the construction the minimal and the maximal extreme positions of the center of mass are relevant. Depending on the variable magnitudes of the partial masses and their lever arms a minimal and a maximal lever arm for the center of mass results. In the field the pilot determines the position and the magnitude of the masses, whereupon the pilot needs to consider that the maximal plane weight mustn't exceed 820 kg. ^[9]

In the following two cases are listed for the longest and shortest lever arm for the center of mass.

Case 1

A maximal lever arm length for the center of mass of XF-max= 525.9 mm results, when just one person flights with a body weight of 70 kg and sits 300 mm backwards the B.P., some luggage is loaded with a weight of 25 kg and with a lever arm of 880 mm and the minimal fuel of 7.7 kg (equal to 10.5 l) with an lever arm of 330 mm is loaded. The tare weight of the plane is assumed with 600 kg and with the most unfavorable lever arm of 540 mm. The partial masses and their lever arms are descriptively shown in table 2.

Maximum lever arm length XF-max: 525.9 mm for:

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Table 2: Set-Up for Xf-max

Case 2

The aircraft is filled-up with 27.4 l that is equal to 20 kg, with a lever arm of 330 mm. The crew consists of two people with a total body weight of 220 kg and they push the seat totally to the front, so that a lever arm of 175 mm is triggered. Due to the fact that it is only a sightseeing flight the crew disclaims any luggage. The tare weight of the plane is still 600 kg but now the other possible unfavorable lever arm of 485 mm is chosen for this calculation. Table 3 shows descriptively the partial masses and their lever arms for the minimal lever arm of XF-min = 400.1 mm.

Minimal lever arm length XF-m;n: 400.1 mm for:

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Table 3: Set-Up for Xf-min

3.1.2 Determination of the horizontal landing gear position under disregard of plane inclination

For first assessment of the horizontal landing gear position Lh, the reference line HE0 in Figure 8 is assumed as horizontal and for this the minimal tail load is considered.

illustration not visible in this excerpt

Figure 8: Reference line HE0^[10]

To handle the sailplane properly on ground the load on the tail wheel shouldn't be too high. Otherwise, the tail load shouldn't be too low, elsewise the plane could overturn. For further calculation the tail load is assumed as 60 kg.

With these allegations it is possible to define the horizontal landing gear position Lh. The value of this position is calculated in equation (4) with moment equilibrium around the landing gear, at which only, due to the simplification, the horizontal distances are considered. The position is dependent of the center of mass position Xf. For that it is considered for the maximal position XF-max and for the minimal position XF-min and its corresponding total plane weight. The length from the tail to the reference line B.P. Ls (cf. Figure 7) is stated with:

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Table 4: Resultant horizontal position Lh for the landing gear

The horizontal position Lh = -58 mm for the minimal center of mass XF-min is chosen. A rearward migrating center of mass will cause a higher tail load. Otherwise, if the position Lh = -129 mm is chosen and in case of a forward migrating center of mass, the tail load will decline and the danger of “overturning” rises.

To determine the vertical position from the fuselage basement to the ground, the propeller clearance needs to be considered. The minimal distance between ground and propeller is stated with 230 mm^[11]. That assumption applies for a horizontal position of HE0. The distance from ground to the reference line HE0 dG-HE0 is determined in equation (6) with the known propeller diameter of 1 600 mm^[12].

The necessary distance between ground and fuselage bottom is called dG-FB. Whereat a simplification is made, it is assumed that the fuselage bottom in this area isn't curved, the area is assumed as flat. Then out of the component drawings the distance between the axle “HE0” and the bottom of the fuselage is stated with 525 mm. The distance dG-FB is calculated in equation (7).

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Consequentially a minimum distance between the ground and the fuselage of 505 mm is necessary.

The position of the landing gear is now roughly defined. The position refers to the contact point of the wheel to the reference line B.P. (horizontal position) and the fuselage bottom (vertical position).

With a further assumption it can said that the tire is flat, and the contact point with the ground is made with the rim. This is assumed to include the compression of the wheel and to generate a higher factor of safety for the case when the tire is flat. Furthermore less damage of the propeller or the fuselage is expected if such a case occurs.

3.1.3 Comparison with available space

In this assessment the position of the retracted and extended carriage should be determined by consideration of the existing flight-relevant parts and by consideration of the possible position of the mounting of the gear swing. For that contemplation it is required that no flight- relevant parts are affected by the landing gear and that at least material as possible should cut out of the fuselage for implementation of the construction.

Position of the wheel under consideration of the steering elements

The wheel, in its retracted position, mustn't cut or touch steering relevant elements of the aircraft. Otherwise, the landing gear swing should be as short as possible to save weight and to get a compact construction. Therefore, the wheel is put as close as possible to the steering rod (Figure 9). The outcoming horizontal position of the wheel center is 725 mm in forefront of the holm. The vertical height in reference to the horizontal console wall part is 360 mm.

The wheel should look out of the fuselage, to raise the safety during a crash landing if the landing gear can't be extended. For that case it will be possible to land the plane anyway without bigger damages.

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Figure 9: Position of the wheel

Position of the landing gear mounting under consideration of the seat frame and holm position

The position of the landing gear mounting depends first on the execution of the mounting itself. It is mounted with a “wall” made of glass fiber layers, which are glued with further layers between the holm and the seat frame. This restricts the horizontal and vertical position of the mounting.

Secondly the mounting should be close at the bottom of the fuselage, the angle of the connection “mounting - center of the wheel hub” and of the fuselage bottom should be zero degrees, so that the gear swing arm closes the fuselage in retracted position. This is made to avoid an additional cover on the gear swing to close the fuselage.

Therefore the best position is 135 mm in forefront of the holm and 360 mm below the horizontal part of the console wall, with a mounting diameter of 50 mm (Figure 10).

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Figure 10: Position of landing gear mounting

3.1.4 Adaption of the landing gear position

With the position stated in the chapter before, the midpoint of the wheel moves on a radius of 590 mm around the mounting. Therefore results the horizontal position Lh of (-) 110 mm to the right of the reference line and the vertical position of 505 mm which is fixed by the JAR 22. In Figure 11 the final position of the wheel in retracted and extended position is shown.

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Figure 11: Final position of the wheel

Due to the new horizontal position of the landing gear in extended position another load at the tail skid results.

Furthermore, the displacement of the center of gravity needs to be considered. If the sailplane stands on the main wheel and on the tail skid wheel; the reference line HEO isn't any more horizontal, it is inclined. That causes a higher tail skid load as in the case there the position of HE0 is determined as horizontal . The length from the reference line HE0 to the contact point of ground with tail wheel is assumed with 400 mm.

The following equations refer to the Figure 7 and Figure 12.The length a, determined in equation (8) and b, determined in equation (9), depends on the position of the center of mass.

The maximal tail load results for the maximal lever arm of the center of mass: XF-max = 525.9 mm and a corresponding weight of 702.67 kg.

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With moment equilibrium around the main gear the tail load can be determined. The forces are acting in vertical direction and the corresponding horizontal lengths are determined with the help of trigonometric functions (cf. equation (10) - (15)).

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Figure 12: Geometry of plane on ground

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In equation (16) the moment equilibrium around the contact point of the main gear wheel and the ground is already converted so that the tail load is directly determined.

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The maximal tail load in the most unfavorable condition is 82.0 kg. In the following chapter the tail loads for different conditions are contemplated.

3.1.5 Provision for the start-kick

To avoid overturning at the most awkward state, the condition at the start needs to be considered. At the start the plane-engine produces a force in horizontal direction that causes a load removal at the tail wheel. These load removal mustn't exceed 82.0 kg, otherwise the plane would overturn and the propeller would be damaged.

This case is described by Figure 13, where Fst is the start kick force by the motor, Mtotal is the total weight forcing in the center of mass, and Ph and Ps are the ground contact forces.

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Figure 13: Start kick

A realistic value for the force Fst is about 1000 N^[13], and the state is determined for the center of mass position with the smallest possible lever arm to the main gear.

In equation (17) a moment equilibrium around the contact point of the main wheel with the ground is postulated to determine the load at the tail skid.

illustration not visible in this excerpt

The tail load Ps in empty condition is 64.2 kg.

Maximal length of the tail-skid set up

The assumption for the tail-skid wheel set up length was minimal 400 mm relating to fuselage to ground. In some cases it could be necessary to swap the tail-skid wheel and replace it by another wheel with different size. For that case it should be examined if the plane turns over, if a bigger sized wheel is used. The situation against overturning is considered like above, when a start-kick of 1 000 N is pulling in horizontal direction.

The maximal length the tail-skid set up can reach is 1 030 mm, in that condition the axle “HE0” is in horizontal position (α = 0°) and the minimal distance between ground and propeller of 230 mm is reached. According to the JAR-22 this distance mustn't fall below 230 mm.^[14] In equation (24) the resultant force Ps at the tail is determined, with the aid of moment equilibrium around the contact point of wheel and ground.

Due to the fact that the angle α is zero, the auxiliary quantity a* is the same like a. For this consideration, the center of mass with the smaller lever arm XF-min is taken and a corresponding weight of 820 kg.

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The result out of this consideration shows that the “tail-skid set up length” can lie between 400 mm and 1 030 mm, with even a high factor of safety avoiding further outside influences which would support overturning.

Lift up of the tail-skid set up

Mere calculative it would be interesting to see how much the tail-skid can be lifted, for e.g. repair, when the airplane is totally empty. The auxiliary quantities result for the center of mass of the empty plane with the smallest lever arm Xl = 485 mm and the corresponding tare weight of Gl = 600 kg. The resulting auxiliary quantities are determined in equation (25) and (26).

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The plane will start overturning, when the center of mass will exceed the horizontal position of the contact point of main wheel and ground, this coherence is depicted in equation (27). This point can be calculated with trigonometric functions in dependence of the angle β (equation (28)) which encases the horizontal line with the inclined reference line HE0. The inclined plane geometry is illustrated in Figure 14.

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Figure 14: Maximal Tail-Skid Length

The distance from the ground to the tail-skid wheel Lmax is determined in equation (29) by adding the vertical components of the distances dG-HE0, a, b and by subtracting the vertical component of the tail wheel length dTS.

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The tail-skid can be maximal 1.57 m lifted till the plane starts overturning. That should be considered if the plane is repaired and the plane rear is therefore raised. However it is possible to load the plane in such a way so that the center of mass postpones in backward direction. This can be achieved for e.g. when the trunk of the plane is loaded, because the

“trunk center of mass” lies much more behind the center of mass of the empty plane. This causes a shift of the whole center of mass in backward direction and the tail-skid can be lifted higher.

4. Load assumption out of the Join Aviation Requirements 22

With the final position of the extended and retracted wheel it is possible to construe the landing gear. The intricacies of the single components can only be construed, by knowing the forces which are acting and causing the different cases of damage.

These load assumptions are already generally made in the JAR 22. In the following these general assumptions are listed and the precise loadings are determined with the help of the known characteristics of the “Valentin Taifun 17E”. The restrictions which are claimed in the JAR 22 are mentioned because they also affect the construction process.

4.1 Level landing condition

Hereinafter the requirements from the JAR 22 are stated which are important for the loads which emerge during the landing.^[15]

Thereout follows that the carriage needs to gather the landing hit without reaching the total impact of the damping elements.^[16]

The magnitude of the kinetic energy is determined by assuming the biggest weight of the plane Mtotal-max and with a descent rate of vd = 1.5 m/s.^[17] At that the acceleration of the center of mass mustn't exceed 4 times the force of gravity.^[18]

These assumptions are summarized in the equation (30). Due to the fact that the carriage should be constructed as easy as possible, it is relinquished on damping elements. The whole energy will be absorbed and transferred by the main wheel.

illustration not visible in this excerpt

The potential energy Epot is dependent from the spring constant and of the range of spring. The spring constant is determined in experiment which is explained and defined in the chapter 6.1.

The resulting load from this contemplation is called Pv. Furthermore it can be assumed that the wing lift is balancing the weight of the sailplane during the landing impact. So the static load doesn't influences the landing force.^[19]

Furthermore, for calculation beside the force Pv the horizontal force Ph needs to be considered. The horizontal load must be combined with the vertical load Pv so that the resultant load acts at an angle at 30° with the vertical line.^[20]

4.2 Lateral load conditions

Additionally it is claimed that it is necessary to contemplate also a lateral load. This load is acting at the contact point of the wheel with the ground.^[21]

The applied lateral load Pside correlates with Pv and if it is used for calculation, another auxiliary vertical quantity Pvertical needs to be considered. The coherence of Pside with Pv is stated in equation (32) and the coherence of Pvertical is stated in equation (33).

illustration not visible in this excerpt

4.3 Lever control

The pilot, when strapped in the seat,^[22] must be able to move each control lever full and without restrictions. Even than when he is wearing winter clothing. It is also necessary that the pilot must reach all controls from his seat, for the case when he is solo flying.

4.4 Limit maneuvering load factor

The limit maneuvering load factor describes the centrifugal force which acts on the plane while flight. These centrifugal forces can become due to acrobatic flights, air holes or other turbulences quit high. The limit maneuvering load factor for an utility plane is set with +5.3 g. 23 At this maximal load factor the landing gear mustn't extend, also like other parts of the plane.^[23]

Furthermore for retractable landing gears it must be shown that extension and retraction of the landing gear are possible without difficulty up to Vlo.^[24]

illustration not visible in this excerpt

4.5 Wheel

The landing gear must belong^[25] to an approved type. The tire must withstand the occurring limit load rating, which is stated in chapter 6.1 and in the following construction phase.

4.6 Factors of Safety

The strength requirements are specified for limit loads and for ultimate loads. The limit loads are defined as the maximal occurring loads during flight/landing, where no yielding occurs. Ultimate loads are multiplied by the factor of safety and are defined where failure occurs.^[26]

The ultimate load is, unless otherwise provided, multiplied by the factor of safety Su = 1.5.^[27]

In some cases it is provided to multiply the factor of safety with another one. This happens for:

a) Castings, with an additional factor of Safety of Sc = 2.0^[28]

b) Bolt or pinned joints, for bearing stress, with an additional factor of safety of Sb = 2.0^[29]

5. Construction possibilities

The construction process began with an investigation of existent carriages and retraction systems. Due to the fact that these carriages were specially adapted to the aircraft's the investigation was not quit insightful. The “Valentin Taifun 17E” needs a carriage fitted to its unique requirements. A list with the important requirements and additional wishes was established:

Requirement specification

- Removing as least as possible existing parts
- No cutting of the steering linkage, pedal linkage and of the seats
- The rim (airless tire) should lie on the previous calculated point of support (ground)
- As little as possible comfort limitation for the passengers
- Half of the rim should look out of the fuselage, for the case that the gear can't be extended
- No damping elements should be implemented

Wishes

- Extension and retraction mechanism should be realized with a hand lever
- As easy as possible in respect of the construction, assembling and maintenance
- As few as possible parts should be used
- fiber materials should be preferred
- no space restriction of the cockpit through off-standing parts of the landing gear o As least as possible comfort limitation for the passengers

Based on this list and the market research it was possible to start the creative phase, where different constructions were sketched and the retraction kinematics were evaluated.

In the following three construction possibilities are present and evaluated; each construction represents a milestone of the creative phase. That means that after each milestone further sketches and considerations were made to improve the previous milestone/construction. The last construction is the one which is elaborated.

5.1 Version 1 - Description and Evaluation

Components

The V1 landing gear consists of following components (Figure 15):

- gear swing
- toggle strut 1 and 2
- gear swing mounting (wall between holm seat frame)
- toggle lever
- hand lever
- connecting components

Each component has different requirements which result from different force transmissions and its corresponding cases of damage.

The gear swing and the holm mounting will absorb the bulk of the occurring landing forces. The strut swing is mainly loaded by pressure in its perpendicular direction and bending, it also closes the fuselage, thereby no part of the strut swing looks out of the fuselage in the retracted state. The fixation of the strut swing is executed with a thorough hollow axle, which connects the swing with the fuselage.

The gear swing mounting will be executed as a composite fiber plate and is glued with composite fiber layers onto the holm and the seat frame, whereof the forces are remitted into the fuselage. The cases of damage which are relevant for the holm mounting plate are

- Failure through bearing stress
- Failure of the adhesive bond
- Bending Failure

The toggle struts are mainly responsible for the extension and retraction mechanism, but also transfers a slight part of the landing forces, in terms of pressure load, into the fuselage. The toggle strut 2 needs to be designed like a fork, which enlaces around the wheel and absorbs the landing forces from the wheel axle. The toggle strut 1 needs to be enlarged to achieve a flushing closure of the fuselage, to avoid an aerodynamically inducement of the airplane. It is also necessary to add an additional arm to the toggle strut 1 for the mounting of the toggle lever.

The toggle lever can be contemplated as a tension rod, in extended situation, and also as compression strut, in retracted situation. The toggle lever is pin jointed with the toggle strut arm and the hand lever. It transmits the movement of the hand lever into the movement of the toggle strut, which causes the extension and the retraction.

The hand lever needs to be ergonomically designed because of high actuating force and to guarantee a high amount of comfort during the actuation. During the actuation the hand lever is loaded by bending, also in the retracted and extended situation, because the position needs to be fixed by a stopper in the console wall.

Retraction and extension mechanism

The extension of the landing gear is supported by the weight and the aerodynamically force of the landing gear. Through a convenient geometry of the struts and levers, an over-twisting of the toggle struts occurs (Figure 16), which causes in the totally extended situation a self­locking effect by the hand lever, whichever is locked by a stopper in the console wall. During the landing the toggle strut will be pushed anti-clockwise and therefore the hand lever would turn clockwise, if it would not be fixed by a stopper (Figure 17). An inadvertent retraction during the landing can be avoided by this “toggle lever mechanism”.

To retract the gear, a support in term of e.g. a gas prop or tension spring is needed. This support has to antagonize the weight and aerodynamically forces of the gear/wheel. It would be indicated that the support has its maximal force in the retracted situation, so that the hand lever automatically turns the system into the retracted position. Furthermore it is needed to fix the position of the hand lever, thereby the gear cannot release out of its position whereas the maximal maneuvering loads acting on it while flight.

Abbildung in dieser Leseprobe nicht enthalten

Figure 15: V1 Retracted position

Abbildung in dieser Leseprobe nicht enthalten

Figure 16: V1 Over-twisting of the toggle struts

Abbildung in dieser Leseprobe nicht enthalten

Figure 17: V1 Extended position

Advantages

- Convenient and safe handling of retraction/extension mechanism through the “toggle lever mechanism”
- Hand lever does not affect the cockpit space

Disadvantages

- Too many parts
- Too many connections with the fuselage » too much cuttings of the fuselage
- Complicated kinematic » a mistake or inaccuracy in the assembly can influence the workability
- Manufacturing and assembly are too complicated o The toggle lever mechanism only works oblique » many linking mountings raising the friction and so the actuation force

[...]

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^[1] Taylor und Munson (1985); p. 744

^[2] El Grafo (2004)

^[3] Carlsson and Kardomateas (2011): p. 7-8

^[4] Valentin GmbH (1983): p. 16

^[5] Valentin GmbH (1983): p. 22

^[6] Valentin GmbH (1987): p. 30

^[7] Joint Aviation Authorities (2001): JAR 22.25 (2)

^[8] Joint Aviation Authorities (2001): JAR 22.925

^[9] Valentin GmbH (1988): p. 9

^[10] Assumption by Prof. Martin Hansen

^[11] Joint Aviation Authorities (2001): JAR 22.925

^[12] Joint Aviation Authorities (2001): JAR 22.725

^[13] Joint Aviation Authorities (2001): JAR 22.725 a)

^[14] Joint Aviation Authorities (2001): JAR 22.725 b)

^[15] Joint Aviation Authorities (2001): JAR 22.725 c)

^[16] Joint Aviation Authorities (2001): JAR 22.473 c)

^[17] Joint Aviation Authorities (2001): JAR 22.479 c)

^[18] Joint Aviation Authorities (2001): JAR 22.473

^[19] Joint Aviation Authorities (2001): JAR 22.725 a)

^[20] Joint Aviation Authorities (2001): JAR 22.725 b)

^[21] Joint Aviation Authorities (2001): JAR 22.725 c)

^[22] Joint Aviation Authorities (2001): JAR 22.777 b)

^[23] Joint Aviation Authorities (2001): JAR 22.337

^[24] Joint Aviation Authorities (2001): JAR 22.729

^[25] Joint Aviation Authorities (2001): JAR 22.731

^[26] Joint Aviation Authorities (2001): JAR 22.301

^[27] Joint Aviation Authorities (2001): JAR 22.303

^[28] Joint Aviation Authorities (2001): JAR 22.621

^[29] Joint Aviation Authorities (2001): JAR 22.623