Patent Description:
There is a concern with respect to the occurrence of corrosion of an aluminum fuselage frame with an aluminum frame being attached to a composite fiber reinforced polymer ("CFRP") skin of an aircraft. In addition, with constructing an aircraft with utilizing, for example, an aluminum frame with a CFRP skin, fatigue issues arise due to the material differences and their different thermal expansion effects. Aluminum, for example, has a thermal expansion co-efficient greater than that of the CFRP. Aluminum experiences contraction in cold conditions and expands in hot conditions relative to the CFRP material which in comparison is thermally neutral. As a result of using different materials for the frame and skin structures, these structures experience compression and tension forces with the aircraft experiencing differing temperatures in operation of the aircraft. Moreover, with respect to production, metallic frames, such as aluminum fuselage frames, require de-burring after drilling to prevent fatigue cracking; require fay surface sealing prior to installation to prevent corrosion; and impose expanded time demands in production.

As a result, there is a need to reduce costs in aircraft production associated with corrosion protection measures being taken with use of different materials such as aluminum for frames and CFRP for skin of the aircraft and to reduce thermal fatigue as a result of thermal expansion effects on interconnected structures in the aircraft which are constructed of different materials each having a different thermal expansion co-efficient. Also, there is a need to shorten production time with respect to the time imparted to production with use of metallic frames.

Document <CIT>, according to its abstract, states fittings for fixing the vertical tail stabilizer of an aircraft in an area of its rear fuselage integrally manufactured with a composite material comprising: a) a first piece comprising lugs for fixing the vertical tail stabilizer and vertical walls for fixing the fitting to the frames; b) at least one pair of additional pieces comprising horizontal walls for fixing the fitting to the skin. The fitting for fixing with an inclined load also comprises a second pair of pieces having an angular shape comprising vertical walls for the fixing with the lugs. The invention also relates to processes for assembling these fittings.

Document <CIT>, according to its abstract, states a frame assembly for a rear section of an aircraft, comprising at least one frame having a plane of symmetry, wherein the frame assembly also comprises at least one supporting element having two ends, wherein each one of the ends is attached to said at least one frame at different sides of the plane of symmetry.

There is provided an aircraft, comprising: a vertical tail fin assembly; a first fuselage frame; and a securement assembly for securing the vertical tail fin assembly to the aircraft, the securement assembly comprising: a first lug member secured to the vertical tail fin assembly; and a first clevis member, wherein: a first end portion of the first clevis member is engaged to the first lug member; a second end portion of the first clevis member is secured to the first fuselage frame constructed of a composite material, with a first fastener which extends through the second end portion and the first fuselage frame in a first direction transverse to the first fuselage frame; the first clevis member has a first prong which defines a first opening at the first end portion of the first clevis member and has a second prong which defines a second opening at the first end portion of the first clevis member; the first lug member is positioned between the first prong and the second prong of the first clevis member; and the securement assembly further includes a first pin which extends through the first opening of the first prong, the second opening of the second prong and through a first lug opening defined by and through the first lug member engaging the first clevis member to the first lug member such that the first pin extends in the first direction transverse to the first fuselage frame. There is further provided a method for securing a vertical tail fin assembly to an aircraft, comprising: securing a first lug member, which is secured to the vertical tail fin assembly, to a first end portion of a first clevis member; and securing a second end portion of the first clevis member to a first fuselage frame constructed of a composite material, with a first fastener which extends through the second end portion of the first clevis member and the first fuselage frame in a first direction transverse to the first fuselage frame; wherein the securing of the first lug member to the first end portion of the first clevis member includes: the first clevis member having a first prong which defines a first opening at the first end portion of the first clevis member and has a second prong which defines a second opening at the first end portion of the first clevis member; the first lug member is positioned between the first prong and the second prong of the first clevis member and defines a first lug opening through the first lug member; and a first pin extends through the first opening of the first prong, the second opening of the second prong and through the first lug opening of the first lug member, such that the first pin extends in the first direction transverse to the first fuselage frame.

The features, functions, and advantages that have been discussed can be achieved independently in various embodiments or may be combined in yet other embodiments as long as these fall within coverage of the appended claims, further details of which embodiments can be seen with reference to the following description and drawings.

In referring to <FIG>, aircraft <NUM> includes fuselage assembly <NUM>, wing assemblies <NUM> and vertical tail fin assembly <NUM>. In fabrication of aircraft <NUM>, vertical tail fin assembly <NUM> includes a support structure (not shown) of spars and ribs to which skin <NUM> of aircraft <NUM> secures. Vertical tail fin assembly <NUM> further secures to fuselage assembly <NUM> of aircraft <NUM> with being connected to fuselage frames. In construction of aircraft <NUM>, as discussed earlier, fabricators have utilized metal, such as aluminum, to construct fuselage frames. Metallic fuselage frames were similar to the configurations of fuselage frames, such as for example, first, second, third and fourth fuselage frames, 20a, 20b, 20c and 20d, as seen in <FIG>. The number of fuselage frames to which vertical tail fin assembly <NUM> is secured can vary depending on the design of the particular aircraft <NUM>.

In the present disclosure, first through fourth fuselage frames 20a-20d, are now constructed with CFRP material, instead of metal, which avoids drawbacks, as discussed earlier associated with metallic fuselage frames. With utilizing CRFP for constructing fuselage frames there is no longer a need to take corrosion resistance preventative measures with respect to a metallic fuselage frame. In addition, with a CFRP constructed fuselage frame there is no longer a need for inspections related to material fatigue with respect to the first through fourth fuselage frames 20a-20d and fuselage skin <NUM>, with first through fourth fuselage frames 20a-20d and skin <NUM> of aircraft <NUM> both being constructed of similar CFRP materials resulting in first through fourth fuselage frames 20a-20d and skin <NUM> having similar thermal expansion co-efficient characteristics. Moreover, extended fabrication times of aircraft <NUM> are reduced with eliminating forging order delays with respect to fabrication of metallic fuselage frames being removed from the fabrication scheduling.

These drawbacks are overcome with first through fourth fuselage frames 20a-20d, in this example, and fuselage skin <NUM> now both being constructed of CFRP material. With the benefits provided with using CFRP material used for fuselage frames such as, 20a-20d, securement assembly <NUM>, to be discussed herein, is needed for securing vertical tail fin assembly <NUM> to aircraft <NUM> through first through fourth fuselage frames 20a-20d of fuselage assembly <NUM>. Securement assembly <NUM> provides needed transferring of shear loads from vertical tail fin assembly <NUM> to first through fourth fuselage frames 20a-20d, of the present example, which are constructed of CFRP material.

Securement of vertical tail fin assembly <NUM> to fuselage frames constructed of metal, was accomplished with bolting of vertical tail fin assembly <NUM> in a generally vertical direction relative to vertical tail fin assembly <NUM> and the metal fuselage frames. However, this configuration of securement, with respect to a fuselage frame now constructed of CFRP material, does not provide optimal shear resistance with respect to shear loadings transmitted from vertical tail fin assembly <NUM> to, in this example, first through fourth fuselage frames 20a-20d as a result of aircraft <NUM> operations. As a result, securement assembly <NUM>, as seen in <FIG>, and described herein, provides an optimal securement of vertical tail fin assembly <NUM> to, in this example, first through fourth fuselage frames 20a-20d constructed of CFRP for resistance to shear loads originating from vertical tail fin assembly <NUM>.

In the present example, of securing vertical tail fin assembly <NUM> to four, first through fourth, fuselage frames 20a-20d herein, each of first through fourth fuselage frames 20a-20d has a pair of securement assemblies <NUM> spaced apart on each of first through fourth fuselage frames 20a-20d securing vertical tail fin assembly <NUM> to each of first through fourth fuselage frames 20a-20d. In referring to <FIG> a starboard side <NUM> of aircraft <NUM> perspective view of first through fourth fuselage frames 20a-20d is seen and in <FIG> a port side <NUM> of aircraft <NUM> perspective view of first through fourth fuselage frames 20a-20d is seen. The two views provide a view of each pair of securement assemblies <NUM> with respect to each of first through fourth fuselage frames 20a-20d. First pair <NUM> of securement assemblies <NUM> are positioned on first fuselage frame 20a, second pair <NUM> of securement assemblies <NUM> are positioned on second fuselage frame 20b, third pair <NUM> of securement assemblies <NUM> are positioned on third fuselage frame 20c and fourth pair <NUM> of securement assemblies <NUM> are positioned on fourth fuselage frame 20d.

In the present example, first and fourth pairs <NUM> and <NUM> of securement assemblies <NUM> are configured the same and are optimal in resisting shear forces transmitted from vertical tail fin assembly <NUM> that have resulting shear loads in forward direction <NUM> and aft direction <NUM>, such as seen in <FIG> and <FIG>. In contrast, in this example, second and third pairs <NUM> and <NUM> of securement assemblies <NUM> are configured the same and are optimal in resisting shear forces transmitted from vertical tail fin assembly <NUM> that have resulting shear loads oriented in starboard direction <NUM> or in port direction <NUM>. The difference between the first pair <NUM> and fourth pair <NUM> of securement assemblies <NUM>, on the one hand, and second pair <NUM> and third pair <NUM> of securement assemblies <NUM>, on the other hand, is the orientation of a pin used in the securement assemblies <NUM> to lug members to be discussed. In first pair <NUM> and fourth pair <NUM> of securement assemblies <NUM> the pin, for each securement assembly <NUM> is positioned to extend along first fuselage frame 20a and fourth fuselage frame 20d, respectively. In second pair <NUM> and third pair <NUM> of securement assemblies <NUM> the pin, for each securement assembly <NUM> is positioned to extend transverse to second fuselage frame 20b and third fuselage frame 20c, respectively.

In the present example, first pair <NUM> of securement assemblies <NUM> is positioned on first fuselage frame 20a and second pair <NUM> of securement assemblies <NUM> is positioned on second fuselage frame 20b as seen in <FIG> and <FIG>. It should be appreciated the location of first fuselage frame 20a carrying first pair <NUM> of securement assemblies <NUM> is not restricted to being positioned as the most aft position <NUM> of first through fourth fuselage frames 20a-20d which secure to vertical tail fin assembly <NUM> and the location of fourth fuselage frame 20d carrying fourth pair <NUM> of securement assemblies <NUM> is not restricted to being positioned on the most forward position <NUM> of first through fourth fuselage frames 20a-20d which secure to vertical tail fin assembly <NUM>. Similarly, second pair <NUM> of securement assemblies <NUM> is not restricted to being positioned on second fuselage frame 20b positioned between first and fourth pairs <NUM>, <NUM> of securement assemblies <NUM>, which are positioned on first and fourth fuselage frames 20a and 20d, respectively. Similarly, third pair <NUM> of securement assemblies <NUM> is not restricted to being positioned on third fuselage frame 20c positioned between the first and fourth pairs <NUM>, <NUM> securement assemblies, which are positioned on first and fourth fuselage frames 20a and 20d, respectively. In the present example described herein, first and fourth pairs <NUM>, <NUM> of securement assemblies <NUM> are positioned on first and fourth fuselage frames 20a and 20d which are positioned most aft position <NUM> and most forward position <NUM> of aircraft <NUM>, respectively, for securing to vertical tail fin assembly <NUM>. Second and third pairs <NUM>, <NUM> are positioned on second and third fuselage frames 20b and 20c, respectively, which are positioned between first and fourth fuselage frames 20a and 20d, for securing vertical tail fin assembly <NUM>.

In referring to <FIG>, securement assembly <NUM> for securing vertical tail fin assembly <NUM> to aircraft <NUM> by way of securing to, in this example, first through fourth fuselage frames 20a-20d, includes first lug member <NUM> is secured (not shown) to vertical tail fin assembly <NUM>, such as for example bolting (not shown) first lug member <NUM> to a framework of spars and ribs (not shown) of vertical tail fin assembly <NUM>. First clevis member <NUM>, positioned on first fuselage frame 20a, has a first end portion <NUM> of first clevis member <NUM> engaged to first lug member <NUM>, as seen in <FIG> and which will be discussed in further detail. First clevis member <NUM> further includes second end portion <NUM> of the first clevis member <NUM> which is secured to first fuselage frame 20a, which is constructed of composite material, CFRP. First fastener <NUM> extends through second end portion <NUM>, as will be further discussed, and first fuselage frame 20a in first direction <NUM> transverse to first fuselage frame 20a. In this example, first fastener <NUM> includes a bolt and nut assembly.

As seen in <FIG>, first clevis member <NUM> has first prong <NUM>, which defines first opening <NUM> at first end portion <NUM> of first clevis member <NUM> and has second prong <NUM> which defines second opening <NUM> at first end portion <NUM> of first clevis member <NUM>. First lug member <NUM> is positioned between first prong <NUM> and second prong <NUM> of first clevis member <NUM>. First pin <NUM> extends through first opening <NUM> of first prong <NUM>, second opening <NUM> of second prong <NUM> and through first lug opening <NUM> defined by and through first lug member <NUM> engaging first clevis member <NUM> to first lug member <NUM> such that first pin <NUM> extends in first direction <NUM> transverse to first fuselage frame 20a. Second end portion <NUM> of first clevis member <NUM> includes first securement flange <NUM> which extends along forward side <NUM> of first fuselage frame 20a with first fastener <NUM> extending through first securement flange <NUM> and first fuselage frame 20a. Positioning first fastener <NUM>, as seen in <FIG> and <FIG>, extending in first direction <NUM> transverse to first fuselage frame 20a, first clevis member <NUM> is secured to first fuselage frame 20a, which is constructed of CFRP, to optimally secure first securement flange <NUM> to first fuselage frame 20a so as to optimally confront shear loadings being received from vertical tail fin assembly <NUM>.

First pair <NUM> of clevis members, as seen in <FIG>, includes first clevis member <NUM> and second clevis member <NUM> secured to first fuselage frame 20a wherein first clevis member <NUM> and second clevis member <NUM> are positioned spaced apart from one another in second direction <NUM> along first fuselage frame 20a. First clevis member <NUM> has first prong <NUM> which defines first opening <NUM> at first end portion <NUM> of first clevis member <NUM> and second clevis member <NUM> has first prong <NUM> which defines first opening <NUM> at first end portion <NUM>, as seen in <FIG>, of second clevis member <NUM> similar to that of first end portion <NUM> of first clevis member <NUM>. First clevis member <NUM> has second prong <NUM> which defines second opening <NUM> at first end portion <NUM> of first clevis member <NUM> and second clevis member <NUM> has second prong <NUM> which defines second opening <NUM>, as seen in <FIG>, at first end portion <NUM>, as shown in <FIG>, of second clevis member <NUM> similar to that of first end portion <NUM> of first clevis member <NUM>, as seen in <FIG>.

First lug member <NUM> is positioned between first prong <NUM> and second prong <NUM> of first clevis member <NUM> of first pair <NUM> of first clevis member <NUM> and second clevis member <NUM>. Second lug member <NUM>, as seen in <FIG> is positioned between first prong <NUM> and second prong <NUM> of <FIG>, of second clevis member <NUM> of first pair <NUM> of first clevis member <NUM> and second clevis member <NUM>. First pin <NUM>, as seen in <FIG>, extends through first opening <NUM> of first prong <NUM> of first clevis member <NUM>, second opening <NUM> of second prong <NUM> of first clevis member <NUM> and through first lug opening <NUM> defined by and through first lug member <NUM> such that first pin <NUM> extends in first direction <NUM> transverse to first fuselage frame 20a. Second pin (not shown) extends through first opening <NUM> of first prong <NUM> of second clevis member <NUM>, second opening <NUM> of second prong <NUM> of second clevis member <NUM> and through second lug opening (not shown) defined by and through second lug member <NUM>, as seen in <FIG>, such that second pin (not shown) extends in first direction <NUM> transverse to first fuselage frame 20a. The configuration of second clevis member <NUM> engaged to second lug member <NUM> is similar as shown in <FIG> with respect to first clevis member <NUM> securing to first lug member <NUM>.

Second end portion <NUM> of first clevis member <NUM> includes, as seen in <FIG>, first securement flange <NUM> and second end portion <NUM> of second clevis member <NUM> includes second securement flange <NUM> as seen in <FIG>. First securement flange <NUM> extends along forward side <NUM> of first fuselage frame 20a with first fastener <NUM>, of <FIG>, extending through first securement flange <NUM> and first fuselage frame 20a in first direction <NUM> transverse to first fuselage frame 20a. Second securement flange <NUM> extends along forward side <NUM> of first fuselage frame 20a with second fastener <NUM>, as seen in <FIG>, extending through second securement flange <NUM> and first fuselage frame 20a in first direction <NUM> transverse to first fuselage frame 20a, similar to the configuration of first securement flange <NUM> as seen in <FIG>.

Third clevis member <NUM>, as seen in <FIG> and <FIG>, is secured to, in this example, to second fuselage frame 20b, constructed of composite material, CFRP, spaced apart from first clevis member <NUM> along length L of aircraft <NUM>, which has first prong <NUM> which defines first opening <NUM> at first end portion <NUM> of third clevis member <NUM> and has a second prong <NUM> which defines second opening <NUM> at first end portion <NUM> of third clevis member <NUM>. Third lug member <NUM> is positioned between first prong <NUM> and second prong <NUM> of third clevis member <NUM>. Third pin <NUM> which extends through first opening <NUM> of first prong <NUM>, second opening <NUM> of second prong <NUM> of third clevis member <NUM> and through third lug opening (not shown) defined by and through third lug member <NUM> engaging third clevis member <NUM> to third lug member <NUM> such that third pin <NUM> extends in third direction <NUM>, as seen in <FIG>, along second fuselage frame 20b. Second end portion <NUM>, as seen in <FIG>, of third clevis member <NUM> includes third securement flange <NUM> which extends along forward side <NUM> of second fuselage frame 20b with third fastener <NUM>, such as for example a bolt with a nut, extending through third securement flange <NUM> and second fuselage frame 20b in fourth direction <NUM> transverse to second fuselage frame 20b.

Further included in this example is second pair of clevis members <NUM>, as seen in <FIG>, including third clevis member <NUM>, as described above, and fourth clevis member <NUM> secured to second fuselage frame 20b, which is also constructed of a composite material, CFRP. Third clevis member <NUM> and fourth clevis member <NUM> are positioned spaced apart from one another in third direction <NUM> along second fuselage frame 20b. As described above third clevis member <NUM> has first prong <NUM> which defines first opening <NUM> at first end portion <NUM> of third clevis member <NUM> and fourth clevis member <NUM> has first prong <NUM> which defines first opening <NUM> at first end portion (not shown) of fourth clevis member <NUM>. First end portion (not shown) of fourth clevis member <NUM> is similar to first end portion <NUM> of third clevis member <NUM> shown in <FIG>. Third clevis member <NUM>, a discussed earlier and seen in <FIG>, has second prong <NUM> which defines second opening <NUM> at first end portion <NUM> of third clevis member <NUM>. Fourth clevis member <NUM> has second prong <NUM> which defines second opening <NUM> at first end portion (not shown) of fourth clevis member <NUM>.

Third lug member <NUM> is positioned between first prong <NUM> and second prong <NUM> of third clevis member <NUM>. Fourth lug member <NUM>, as seen in <FIG>, is positioned between first prong <NUM> and second prong <NUM> of fourth clevis member <NUM> of <FIG>. Third pin <NUM>, as seen in <FIG>, extends through first opening <NUM> of first prong <NUM> of third clevis member <NUM>, second opening <NUM> of second prong <NUM> of third clevis member <NUM> and through third lug opening (not shown) defined by and through third lug member <NUM>, such that third pin <NUM> extends in third direction <NUM>, as shown in <FIG>, along second fuselage frame 20b. Fourth pin (not shown) but similar to third pin <NUM> as seen in <FIG>, extends through first opening <NUM> of first prong <NUM> of fourth clevis member <NUM>, second opening <NUM> of second prong <NUM> of fourth clevis member <NUM>, of <FIG>, and through a fourth lug opening (not shown) defined by and through fourth lug member <NUM>, as seen in <FIG>, such that fourth pin (not shown) extends in third direction <NUM> along second fuselage frame 20b as seen in <FIG>.

Second end portion <NUM> of third clevis member <NUM>, as seen in <FIG>, includes third securement flange <NUM> and second end portion (not shown) of fourth clevis member <NUM> includes fourth securement flange <NUM>, as seen in <FIG>. Second end portion (not shown) of fourth clevis member <NUM> and fourth securement flange <NUM> are similar to second end portion <NUM> of third clevis member <NUM> and third securement flange <NUM> as shown in <FIG>. Third securement flange <NUM> extends along forward side <NUM>, as seen in <FIG> and <FIG>, of second fuselage frame 20b with third fastener <NUM> extending through third securement flange <NUM> and second fuselage frame 20b in fourth direction <NUM> transverse to second fuselage frame 20b. Fourth securement flange <NUM>, as seen in <FIG>, extends along forward side <NUM> of second fuselage frame 20b with fourth fastener (not shown) however similar to that of third fastener <NUM> of <FIG> extending through the fourth securement flange <NUM> and second fuselage frame 20b in fourth direction <NUM> transverse to second fuselage frame 20b, similar to third securement flange <NUM> securement arrangement as seen in <FIG>.

First through fourth fuselage frames 20a-20d, in this example, are constructed with composite material such as CFRP. In the present example, the composite material has a five ply configuration for each layer of composite material used in constructing first through fourth fuselage frames 20a-20d, wherein in this example multiple layers of composite material are used. In referring to <FIG> one ply has a nonlinear fiber configuration <NUM> of fibers. In constructing each layer of composite material, one ply of composite material contains the nonlinear fiber configuration <NUM> of fibers which includes twenty percent (<NUM>%) of the plies for that layer. In this example, nonlinear fiber configuration <NUM> of fibers extend in a curved direction having a radial axis R. Two plies of composite material have a first linear fiber configuration of fibers <NUM> which extend within an angular range which includes plus or minus five degrees of plus thirty degrees of being angularly displaced from radial axis R. The angular displacement is represented as angle "A". Two plies of composite material have first linear fiber configuration of fibers <NUM> which includes about forty percent (<NUM>%) of the plies for that layer. Another two plies of composite material for the layer includes a second linear fiber configuration of fibers <NUM> which extend within an angular range which includes plus or minus five degrees of minus thirty degrees of being angularly displaced from radial axis R. The angular displacement is represented as angle "B". This example of fiber configuration in the composite material provides a construction that is lighter in weight than a metal counterpart fuselage frame and provides required resistance to shear loads originating from vertical tail fin assembly <NUM> as a result of aircraft <NUM> operations.

In referring to <FIG>, a schematic cross section of aircraft <NUM> is shown with fuselage assembly <NUM> and vertical tail fin assembly <NUM> in viewing toward aft position <NUM> of aircraft <NUM>. An example of reactant shear force with respect to first through fourth fuselage frames 20a-20d during operation of aircraft <NUM> is shown. In operation of aircraft <NUM>, an aerodynamic operational force F can be applied, as in this example, from starboard side <NUM> of aircraft <NUM>. The aerodynamic operational force F resultant force load is transmitted through vertical tail fin assembly <NUM> such that fuselage frames such as 20a-20d provide reactant forces of F <NUM> in a direction toward vertical tail fin assembly <NUM> on port side <NUM> of vertical tail fin assembly <NUM> of aircraft <NUM> and of F2 in a direction away from vertical tail fin assembly <NUM> on starboard side <NUM> of vertical tail fin assembly <NUM> of aircraft <NUM> in countering aerodynamic operational force F. With aerodynamic operational force F reversed in direction reactant forces F <NUM> and F2 are reversed in direction. In the example of securement assemblies <NUM> described herein, second and third pairs <NUM> and <NUM> of securement assemblies <NUM> positioned on fuselage frames 20b and 20c, respectively, provide optimal load transfers between fuselage frames 20b and 20c in addressing the aerodynamic operational force F experienced from starboard or port sides <NUM>, <NUM> of aircraft <NUM>. First and fourth pairs <NUM> and <NUM> of securement assemblies <NUM> positioned on fuselage frames 20a and 20d, respectively, provide optimal load transfers between fuselage frames 20a and 20d and vertical tail fin assembly <NUM> in addressing longitudinal operational forces on vertical tail fin assembly <NUM> along length L of aircraft <NUM>.

In referring to <FIG>, method <NUM> for securing vertical tail fin assembly <NUM> to aircraft <NUM> includes securing <NUM> first lug member <NUM>, which is secured to vertical tail fin assembly <NUM>, to first end portion <NUM> of first clevis member <NUM>. Method <NUM> further includes securing <NUM> second end portion <NUM> of first clevis member <NUM> to first fuselage frame 20a constructed of composite material with first fastener <NUM> which extends through second end portion <NUM> of first clevis member <NUM> and first fuselage frame 20a in first direction <NUM> transverse to first fuselage frame 20a.

Securing <NUM> of first lug member <NUM> to first end portion <NUM> of first clevis member <NUM> includes first clevis member <NUM> having first prong <NUM> which defines first opening <NUM> at first end portion <NUM> of first clevis member <NUM> and has second prong <NUM> which defines second opening <NUM> at first end portion <NUM> of first clevis member <NUM>, as seen in <FIG>. First lug member <NUM> is positioned between first prong <NUM> and second prong <NUM> of first clevis member <NUM> and defines first lug opening <NUM> through first lug member <NUM>, as seen in <FIG>. First pin <NUM> extends through first opening <NUM> of first prong <NUM>, second opening <NUM> of second prong <NUM> and through first lug opening <NUM> of first lug member <NUM>, such that first pin <NUM> extends in first direction <NUM> transverse to first fuselage frame 20a.

Method <NUM> for securing further includes first pair <NUM> of clevis members including first clevis member <NUM> and second clevis member <NUM>, as seen in <FIG>, wherein second clevis member <NUM> is secured to first fuselage frame 20a spaced apart from first clevis member <NUM> in second direction <NUM> along first fuselage frame 20a. Second pin (not shown) extends through first opening <NUM> of first prong <NUM> of first portion of the second clevis member <NUM>, second opening <NUM> of second prong <NUM> of first end portion <NUM> of second clevis member <NUM> and through second lug opening (not shown) of second lug member <NUM>, such that second pin (not shown) extends in first direction <NUM> transverse to first fuselage frame 20a. Second fastener <NUM> extends through second end portion <NUM>, as seen in <FIG>, of second clevis member <NUM> and through first fuselage frame 20a in first direction <NUM> transverse to first fuselage frame 20a securing second clevis member <NUM> to first fuselage frame 20a.

Method <NUM> further includes securing of third lug member <NUM>, as seen in <FIG>, to first end portion <NUM> of third clevis member <NUM> with third securement flange <NUM> of a second end portion <NUM> of the third clevis member <NUM> secured to second fuselage frame 20b, which is spaced apart from first fuselage frame 20a. Third clevis member <NUM> includes first prong <NUM> which defines first opening <NUM> at first end portion <NUM> of third clevis member <NUM> and has second prong <NUM> which defines second opening <NUM> at first end portion <NUM> of third clevis member <NUM>, as seen in <FIG>. Third lug member <NUM> is positioned between first prong <NUM> and second prong <NUM> of third clevis member <NUM> and defines third lug opening (not shown) through third lug member <NUM>. Third pin <NUM> extends through first opening <NUM> of first prong <NUM>, second opening <NUM> of second prong <NUM> and through third lug opening (not shown) of third lug member <NUM>, such that third pin <NUM> extends in third direction <NUM> along second fuselage frame 20b. Third fastener <NUM> extends through third securement flange <NUM> of second end portion <NUM> of third clevis member <NUM> and second fuselage frame 20b in fourth direction <NUM> transverse to second fuselage frame 20b, as seen in <FIG>.

Second pair of clevis members <NUM> including third clevis member <NUM> and fourth clevis member <NUM>, as seen in <FIG>, wherein fourth clevis member <NUM> is secured to second fuselage frame 20b spaced apart from third clevis member <NUM> in third direction <NUM> along second fuselage frame 20b. Fourth clevis member <NUM> has first prong <NUM> which defines first opening <NUM> at first end portion (not shown) of fourth clevis member <NUM> and has second prong <NUM> which defines second opening <NUM> at first end portion (not shown) of fourth clevis member <NUM>. First end portion (not shown) of fourth clevis member <NUM> is similar to first end portion <NUM> of third clevis member <NUM> shown in <FIG>. Fourth lug member <NUM>, as seen in <FIG>, is positioned between first prong <NUM> and second prong <NUM> of fourth clevis member <NUM> and defines fourth lug opening (not shown) through fourth lug member <NUM>. Fourth pin (not shown), similar to third pin <NUM> as seen in <FIG>, extends through first opening <NUM> of first prong <NUM>, second opening <NUM> of second prong <NUM> and through fourth lug opening (not shown) of the fourth lug member <NUM>, such that the fourth pin (not shown) extends in third direction <NUM> along second fuselage frame 20b. Fourth fastener (not shown), similar to that of third fastener <NUM> of <FIG>, extends through fourth securement flange <NUM>, as seen in <FIG>, of second end portion (not shown) of fourth clevis member <NUM> and second fuselage frame 20b in fourth direction <NUM> transverse to second fuselage frame 20b, similar to third securement flange <NUM> securement arrangement as seen in <FIG>.

The composite material used in construction of first through fourth fuselage frames 20a-20d include five ply configurations for a layer, wherein multiple layers are employed in this example, as described earlier, which includes one ply having a nonlinear fiber configuration <NUM> of fibers, as seen in <FIG>. Two plies have a first linear fiber configuration of fibers <NUM> which extends within an angular range which includes plus or minus five degrees of plus thirty degrees of angular displacement from radial axis R of nonlinear fiber configuration <NUM> of fibers. Another two plies have a second linear fiber configuration of fibers <NUM> which extends within an angular range which includes plus or minus five degrees of minus thirty degrees of angular displacement from radial axis R of nonlinear fiber configuration <NUM> of fibers.

Claim 1:
An aircraft (<NUM>), comprising:
a vertical tail fin assembly (<NUM>);
a first fuselage frame (20a); and
a securement assembly (<NUM>) for securing the vertical tail fin assembly (<NUM>) to the aircraft (<NUM>), the securement assembly (<NUM>) comprising:
a first lug member (<NUM>) secured to the vertical tail fin assembly; and
a first clevis member (<NUM>), wherein:
a first end portion (<NUM>) of the first clevis member is engaged to the first lug member;
a second end portion (<NUM>) of the first clevis member is secured to the first fuselage frame (20a) constructed of a composite material, with a first fastener (<NUM>) which extends through the second end portion and the first fuselage frame in a first direction transverse to the first fuselage frame;
the first clevis member (<NUM>) has a first prong (<NUM>) which defines a first opening (<NUM>) at the first end portion (<NUM>) of the first clevis member (<NUM>) and has a second prong (<NUM>) which defines a second opening (<NUM>) at the first end portion (<NUM>) of the first clevis member (<NUM>);
the first lug member (<NUM>) is positioned between the first prong (<NUM>) and the second prong (<NUM>) of the first clevis member (<NUM>); and
the securement assembly (<NUM>) further includes a first pin (<NUM>) which extends through the first opening (<NUM>) of the first prong (<NUM>), the second opening (<NUM>) of the second prong (<NUM>) and through a first lug opening (<NUM>) defined by and through the first lug member (<NUM>) engaging the first clevis member (<NUM>) to the first lug member (<NUM>) such that the first pin (<NUM>) extends in the first direction (<NUM>) transverse to the first fuselage frame (20a).