A method of preparing a beta titanium-composite laminate comprising providing a first layer of beta titanium alloy having a yield strength to modulus of elasticity ratio and adhering a first layer of composite having a strength to modulus of elasticity ratio to the layer of beta titanium alloy, thereby forming a beta titanium-composite laminate, where the yield strength to modulus of elasticity ratio of the first layer of beta titanium alloy matches the strength to modulus of elasticity ratio of the first layer of composite such that the first layer of beta titanium alloy will reach its stress limit and the first layer of composite will reach its stress limits at about the same total strain. Also, a beta titanium-composite laminate comprising a first layer of beta titanium alloy having a yield strength to modulus of elasticity ratio, and a first layer of composite having a strength to modulus of elasticity ratio adhered to the first layer of beta titanium alloy, where the yield strength to modulus of elasticity ratio of the first layer of beta titanium alloy matches the strength to modulus of elasticity ratio of the first layer of composite such that the first layer of beta titanium alloy will reach its stress limit and the first layer of composite will reach its stress limits at about the same total strain. Further, a structure such as an airplane part including a beta titanium-composite laminate according to the present invention.

BACKGROUND
 Aircraft primary structures are predominately made from non-composite
 metals. However, the aerospace industry has been increasingly using light
 weight, advanced composite materials in place of metals to produce primary
 structures because of the high specific strength of advanced composites
 materials. Nevertheless, advanced composite materials have not entirely
 replaced metals in primary structures because advanced composites are more
 sensitive to damage, have lower bearing strength, and are more susceptible
 to fastener failure than metals.
 Several improved composites have been designed, including Arall, as
 disclosed in U.S. Pat. No. 4,500,589, and Glair, as disclosed in U.S. Pat.
 No. 5,039,571. Disadvantageously, however, the layers of both Arall and
 Glair have a mismatch of the ratio between their modulus of elasticity and
 their yield strength.
 For example, Arall is a composite of aluminum skins adhesively bonded to a
 core of Aramid fiber/epoxy composite. The Aramid fiber of Arall has a
 unidirectional yield strength of about 172,000 psi and a modulus of
 12.2.times.10.sup.6 psi, while the aluminum layer has a yield strength of
 50,000 psi and modulus of 10.0.times.10.sup.6 psi. Thus, stressing the
 Aramid fiber layer to its maximum yield strength would stress the aluminum
 layer to 141,000 psi, which is well above the maximum limit for the
 aluminum layer. Conversely, stressing the aluminum layer to its maximum
 yield strength of 50,000 psi stresses the Aramid fiber layer to 61,000
 psi, which is well below the maximum limit for the Aramid fiber layer.
 Thus, the strength of the Aramid fiber layer is underutilized. Similarly,
 the layers of composite laminates of standard alpha-beta alloys of
 titanium, such as Ti6Al-4V, and carbon fiber composites have a mismatch of
 the ratio between their modulus of elasticity and their yield strength.
 The aerospace industry has not used the newer beta alloys of titanium, such
 as TIMETAL.RTM. 15-3 (Ti-15V-3Cr-3Sn-3Al) and TIMETAL.RTM. 21s
 (Ti-15Mo-3Al-3Nb), for composite laminates even though these beta alloys
 of titanium have higher strength and a lower modulus of elasticity because
 commonly used adhesives will not stick adequately to the titanium oxide
 surface layer present on these alloys. While methods have been developed
 to bond the standard alpha-beta alloys of titanium to carbon fiber
 composites, these methods do not work with the beta alloys of titanium.
 Therefore, there is a need for improved composite laminates for the primary
 structures of aircraft which utilize the full strength of each layer.
 Further, there is a need for a method of preparing these composite
 laminates.
 SUMMARY
 According to one embodiment of the present invention, there is provided a
 method of preparing a beta titanium-composite laminate. The method
 comprises providing a first layer of beta titanium alloy having a yield
 strength to modulus of elasticity ratio and providing a first layer of
 composite having a yield strength to modulus of elasticity ratio. Then,
 the first layer of beta titanium alloy is adhered to the first layer of
 composite, thereby forming a beta titanium-composite laminate. The yield
 strength to modulus of elasticity ratio of the first layer of beta
 titanium alloy matches the strength to modulus of elasticity ratio of the
 first layer of composite such that the first layer of beta titanium alloy
 will reach its stress limit and the first layer of composite will reach
 its stress limits at about the same total strain.
 In one embodiment, the beta titanium alloy provided is selected from the
 group consisting of (Ti-15V-3Cr-3Sn-3Al) and (Ti-15Mo-3Al-3Nb). In a
 preferred embodiment, the composite provided is a carbon fiber/epoxy
 composite. In another preferred embodiment, the composite provided is an
 S2-glass/epoxy composite.
 In one embodiment, adhering comprises applying an adhesive to the beta
 titanium alloy. In another embodiment, adhering comprises bonding the
 composite by heating the composite.
 In a particularly preferred embodiment, the yield strength to modulus of
 elasticity ratio of the first layer of beta titanium alloy is between
 about 5% of the strength to modulus of elasticity ratio of the first layer
 of composite. In another particularly preferred embodiment, the method
 additionally comprises cold reducing the beta titanium alloy before
 adhering. In yet another particularly preferred embodiment, the method
 additionally comprising heating the beta titanium alloy at a temperature
 for a time to produce an aged beta titanium alloy before adhering, such as
 a temperature of approximately 950.degree. F. and a time of approximately
 8 hours.
 In a particularly preferred embodiment, the method additionally comprises
 cold reducing the beta titanium alloy and then aging the beta titanium
 alloy before adhering. In another particularly preferred embodiment, the
 method additionally comprises coating the surface of the beta titanium
 alloy with a metal selected from the group consisting of platinum and the
 functional equivalent of platinum as a coating material, to produce a
 coated titanium alloy before adhering. In another particularly preferred
 embodiment, the method additionally comprises abrading the surface of the
 beta titanium alloy before adhering.
 In another particularly preferred embodiment, the method additionally
 comprises, after adhering, providing a second layer of beta titanium alloy
 having a yield strength to modulus of elasticity ratio and adhering the
 second layer of beta titanium alloy to the beta titanium-composite
 laminate, where the yield strength to modulus of elasticity ratio of the
 second layer of beta titanium alloy matches the strength to modulus of
 elasticity ratio of the first layer of composite such that the second
 layer of beta titanium alloy will reach its stress limit and the first
 layer of composite will reach its stress limits at about the same total
 strain. In yet another particularly preferred embodiment, the method
 additionally comprises, after adhering, providing a second layer of
 composite having a strength to modulus of elasticity ratio and adhering
 the second layer of composite to the beta titanium-composite laminate,
 where the strength to modulus of elasticity ratio of the second layer of
 composite matches the yield strength to modulus of elasticity ratio of the
 first layer of beta titanium alloy such that the second layer of composite
 will reach its stress limit and the first layer of beta titanium alloy
 will reach its stress limits at about the same total strain.
 The present invention also includes a method of making an airplane part
 comprises preparing a beta titanium-composite laminate according to the
 present invention and incorporating the beta titanium-composite laminate
 into an airplane part. The present invention also includes a method of
 making an airplane comprising preparing an airplane part according to the
 present invention and incorporating the part into an airplane. The
 airplane part can be selected from the group consisting of airplane skin,
 a spar, a plate and a tube.
 In a preferred embodiment, the present invention includes a beta
 titanium-composite laminate produced according to a method of the present
 invention.
 In another preferred embodiment, the present invention includes a beta
 titanium-composite laminate comprising a first layer of beta titanium
 alloy having a yield strength to modulus of elasticity ratio, and a first
 layer of composite having a strength to modulus of elasticity ratio
 adhered to the first layer of beta titanium alloy. The yield strength to
 modulus of elasticity ratio of the first layer of beta titanium alloy
 matches the strength to modulus of elasticity ratio of the first layer of
 composite such that the first layer of beta titanium alloy will reach its
 stress limit and the first layer of composite will reach its stress limits
 at about the same total strain. For example, the yield strength to modulus
 of elasticity ratio of the first layer of beta titanium alloy can be
 between about 5% of the strength to modulus of elasticity ratio of the
 first layer of composite. The beta titanium-composite laminate can
 additionally comprises a layer of platinum between the first layer of beta
 titanium alloy and the first layer of composite.
 In another particularly preferred embodiment, the beta titanium-composite
 laminate can additionally comprising a second layer of beta titanium alloy
 having a yield strength to modulus of elasticity ratio adhered to the
 first layer of composite. The yield strength to modulus of elasticity
 ratio of the second layer of beta titanium alloy matches the strength to
 modulus of elasticity ratio of the first layer of composite such that the
 second layer of beta titanium alloy will reach its stress limit and the
 first layer of composite will reach its stress limits at about the same
 total strain. The beta titanium-composite laminate can additionally
 comprising a layer of platinum between the second layer of beta titanium
 alloy and the first layer of composite.
 In another particularly preferred embodiment, the beta titanium-composite
 laminate can additionally comprises a second layer of composite having a
 yield strength to modulus of elasticity ratio adhered to the first layer
 of beta titanium alloy. The yield strength to modulus of elasticity ratio
 of the second layer of composite matches the strength to modulus of
 elasticity ratio of the first layer of beta titanium alloy such that the
 second layer of composite will reach its stress limit and the first layer
 of beta titanium alloy will reach its stress limits at about the same
 total strain. For example, the beta titanium-composite laminate of claim
 34, where the yield strength to modulus of elasticity ratio of the second
 layer of beta titanium alloy is between about 5% of the strength to
 modulus of elasticity ratio of the first layer of composite. The beta
 titanium-composite laminate can additionally comprising a layer of
 platinum between the first layer of beta titanium alloy and the second
 layer of composite.
 In a particularly preferred embodiment, the present invention includes an
 airplane or an airplane part comprising a beta titanium-composite laminate
 according to the present invention. The present invention further includes
 an airplane comprising an airplane part according to the present
 invention.

DESCRIPTION
 The present invention addresses a need in the aerospace industry for
 composites materials which can be used to reduce the weight of primary
 structures incorporated into aircraft while providing superior strength
 and damage resistance of the structures. In one embodiment, the present
 invention includes a composite laminate of at least one layer of beta
 titanium alloy that is bonded to at least one layer of a high-strength
 carbon (graphite) fiber/epoxy to produce a beta titanium-carbon fiber
 composite laminate. This composite laminate maintains the advantages of
 each material without sacrificing the load carrying ability or either
 layer by matching the strength to modulus of elasticity ratio between the
 at least one beta titanium alloy layer and the at least one carbon fiber
 layer.
 In a preferred embodiment, the beta titanium-carbon fiber composite
 laminate according to the present invention includes at least one beta
 titanium layer that is made of TIMETAL.RTM. 15-3 (Ti-15V-3Cr-3Sn-3Al)
 (available from Titanium Metals Corporation, Denver, Co., US) and the at
 least one carbon fiber/epoxy composite layer that is made of Hercules.RTM.
 carbon fiber AS4. The beta titanium layer of TIMETAL.RTM. 15-3 is cold
 reduced from the solution treated condition by 50% which gives it a
 tensile strength of 173,000 psi and a modulus of elasticity of
 12.4.times.10.sup.6 psi (a yield strength to modulus of elasticity ratio
 of 0.01395). The carbon composite layer has a unidirectional tensile
 strength of 220,000 psi and a modulus of elasticity of 16.times.10.sup.6
 psi (a strength to modulus of elasticity ratio of 0.01375). Thus, the
 yield strength to modulus of elasticity ratio of the beta titanium alloy
 layer matches the strength to modulus of elasticity ratio of the carbon
 composite layer within about 1.5%. Under the same total strain, the beta
 titanium layer will be stressed to 170,000 psi and the carbon layer will
 be stressed to 220,000 psi. Therefore, each layer of the beta
 titanium-carbon fiber composite laminate will reach its stress limit at
 about the same total strain giving the beta titanium-carbon fiber
 composite laminate a superior specific strength without under-utilizing
 either layer.
 The 50% cold reduction of the titanium layer in the above embodiment is
 made, for example, by rolling a 0.020 inches thick strip of titanium alloy
 in the solution treated condition to a thickness of 0.010 inches. Cold
 reductions of greater or lesser percentages can also be made in order to
 match the strength to modulus of elasticity ratio of the titanium alloy to
 the strength to modulus of elasticity ratio of the composite layer.
 In another embodiment, the beta titanium-carbon fiber composite laminate
 according to the present invention includes at least one beta titanium
 layer and at least one carbon fiber/epoxy composite layer of Hercules.RTM.
 fiber IM8. The beta titanium alloy is aged at 950.degree. F. for 8 hours
 to have a tensile strength of 224,000 psi and a modulus of elasticity of
 16.3.times.10.sup.6 psi (a yield strength to modulus of elasticity ratio
 of 0.01374). The at least one layer of Hercules.RTM. IM8 fiber has a
 unidirectional tensile strength of approximately 380,000 psi and a modulus
 of elasticity of 27.times.10.sup.6 psi (a strength to modulus of
 elasticity ratio of 0.01407). Thus, the yield strength to modulus of
 elasticity ratio of the layer of beta titanium alloy matches the strength
 to modulus of elasticity ratio of the layer of carbon composite within
 about 2.4%. Under the same total strain conditions, the beta titanium
 layer will be stressed to 224,000 psi and the carbon layer will be
 stressed to 371,000 psi. Therefore, each layer of the beta titanium-carbon
 fiber composite laminate will reach its stress limit at about the same
 total strain giving the beta titanium-carbon fiber composite laminate a
 superior specific strength without underutilizing either layer.
 Composite laminates of the present invention can also be made using other
 materials such as S.sub.2 -Glass and boron instead of carbon composite.
 For example, boron in compression has a strength of 400,000 psi and a
 modulus of 30.times.10.sup.6 psi (a strength to modulus of elasticity
 ratio of 0.01333). When boron is combined with at least one layer of cold
 reduced TIMETAL.RTM. 15-3, similar to the composite laminates disclosed
 above, stressing the boron layer to 400,000 psi will stress the titanium
 layer to 217,000 psi. Thus, the yield strength to modulus of elasticity
 ratio of the layer of beta titanium alloy matches the strength to modulus
 of elasticity ratio of the layer of boron within about 4.6%.
 Further, composites of the present invention can also be made using other
 materials such as aluminum or steel alloys instead of beta titanium
 alloys. However, beta titanium alloys are preferred in composites of the
 present invention because of the high specific strength of the beta
 titanium alloy and carbon fiber composites. For example, in an
 aluminum-carbon fiber composite laminate composed of 7075-T6 aluminum
 alloy and high-modulus carbon fiber/epoxy composite, the carbon layer will
 have a tensile strength of 122,000 psi and a modulus of
 27.5.times.10.sup.6 and will stress the aluminum layer to 46,000 psi,
 which is less than the stress limit of the aluminum layer. Therefore, each
 layer of the aluminum-carbon fiber composite laminate will not reach its
 stress limit at about the same total strain.
 Similarly, in a steel-carbon fiber composite laminate of 301 stainless
 steel and high-modulus carbon fiber/epoxy composite, the carbon layer will
 have a tensile strength of 122,000 psi and a modulus of
 27.5.times.10.sup.6. At the maximum tensile strength of the carbon layer,
 the steel layer will stress to 133,000 psi, which is less than the stress
 limit of 301 stainless steel. Therefore, each layer of the steel-carbon
 fiber composite laminate will not reach its stress limit at about the same
 total strain.
 According to one embodiment of the present invention, there is provided a
 method of making a beta titanium-carbon fiber composite laminate. In one
 embodiment, the method includes forming a basic shape from a titanium
 alloy such as TIMETAL.RTM. 15-3 in the solution treated condition. The
 shape is then cleaned by placing the titanium in a hot caustic solution
 followed by a hydrofluoric/nitric acid pickle.
 Next, a surface that is to be bonded to a composite layer is sandblasted
 with an aluminum oxide grit 100 mesh under approximately 40 psi of air
 pressure. This sandblasting creates an irregular surface with a surface
 area that is increased by approximately 100% by the sandblasting.
 The surface is then electrolytically coated with a metallic coating such as
 a thin layer of platinum or a functional equivalent material such as
 nickel. Next, the shape is aged to its final strength which simultaneously
 bonds the coating to the surface of the beta titanium. Aging the coated
 titanium causes the surface oxide of the titanium to go into solution,
 which in turn allows the surface coating to chemically bond to the
 titanium causing an alloy of titanium and platinum to form at the
 interface. The time and temperature of the aging process is selected by
 empirical testing to produce a beta titanium layer which matching the
 strength to modulus of elasticity ratio between the beta titanium alloy
 layer and the carbon fiber layer, while allowing the surface of the beta
 titanium to partly absorb the surface coating. For example, the titanium
 can be aged at approximately 950.degree. F. for 8 hours.
 Next, the shapes are cleaned and primed with a standard primer such as BAR
 127, and then the carbon composite layer is adhesively bonded to the
 primed side of the beta titanium layer with an adhesive such as AF163-2.
 The surfaces are selectively strengthened by placing carbon fabric or
 unidirectional pregreg tape in an amount and orientation that will yield
 the required strength needed for the part being produced without adding
 unnecessary weight. The part is then hot cured in a standard way either in
 a press or autoclave. The resulting part has a much higher specific
 strength and, thus, a lower weight for a given strength than a part
 produced by prior art methods.
 Referring now to FIG. 1, there is shown a chart comparing the yield
 strength, density and specific strength for materials known in the prior
 art with a five layer titanium-carbon fiber composite laminate according
 to the present invention having three layers of 0.010 inches thick
 TIMETAL.RTM. 15-3 and two layers of 0.030 inches thick unidirectional
 Hercules.RTM. fiber IM8. As can be seen, the specific strength of the beta
 titanium-carbon fiber composite laminate of the present invention is
 significantly greater than the prior art materials.
 The specific strength of an all-carbon composite structure can be even
 higher than the specific strength of a titanium-carbon fiber composite
 laminate. However, all-carbon fiber composites are disadvantageously more
 sensitive to damage compared to titanium-carbon fiber composite laminates.
 For example, referring now to FIG. 2, there is shown a chart comparing the
 damage sensitivity of an all-carbon fiber composite (AS4 3501-6
 Quasi-Isotropic), top, with a titanium-carbon fiber composite laminate
 according to the present invention, bottom, in two tests of damage
 sensitivity. As can be seen, the titanium-carbon fiber composite laminate
 was approximately twice as strong as the all-carbon fiber composite in
 both tests.
 To make a part with a beta titanium layer adhered to a carbon composite,
 according to the present invention and as shown in FIG. 3, the titanium
 has to be prepared so that the adhesive will effectively bond to the
 titanium. Normally, the titanium oxide coating on the surface of beta
 titanium alloy will not bond effectively to an adhesive. This is evidenced
 by the industry standard wedge crack test. In this test, a wedge is
 inserted between the strips of titanium, cracking the adhesive. The strips
 of titanium, with the inserted wedge and cracked adhesive, are then placed
 in a humidity chamber. This leaves a strain on the remaining
 adhesive/titanium bond line under an elevated temperature and high
 humidity conditions (140.degree. F., 95% relative humidity). Using the
 industry standard surface preparation methods, such as anodizing, the
 adhesive will completely slick off beta titanium alloys.
 Referring now to FIG. 3, there is shown a diagram of a cross-sectional view
 of a titanium/adhesive-carbon bond according to the present invention
 after the carbon has been bonded and cured with the titanium. The beta
 titanium alloy layer 50 has been sandblasted and then, coated with a
 platinum layer 54 on one side. A standard surface treatment of the carbon
 composite 60 can be used when bonding or co-curing to the primered
 titanium surface. During the aging process, the titanium becomes
 chemically active and starts to absorb the platinum which creates a
 titanium platinum alloy layer 52 that serves to bond the platinum coating
 54 to the titanium layer 50. Then, the platinum layer 54 is coated with a
 primer layer 56, which in turn is coated with adhesive layer 58. Next, the
 carbon layer 60 is bonded to the adhesive layer 58 in the standard manner,
 such as heating the carbon layer 60 and adhesive layer 58 to 350.degree.
 F. for approximately one hour. Information on suitable temperatures and
 pressures for each adhesive and carbon fiber composite are generally
 available through the manufacturer of the carbon fiber composites.
 Aircraft parts according to the present invention can be made having
 suitable strength and increased damage tolerance while weighting
 approximately 50-75% less. For example, an airplane skin made of a
 titanium-carbon fiber composite laminate according to the present
 invention, will have a reduced tendency to tear apart in the event of an
 explosion or gun shot through the skin. Additionally, beta titanium-carbon
 fiber composite laminate sheets according to the present invention have a
 stronger joint than sheets of pure carbon epoxy in which fasteners tend to
 pull through the carbon. See, for example, FIG. 2.
 The methods and composites of the present invention can be used to make
 several parts used in aircraft, including spars, plates and tubes. For
 example, referring now to FIG. 4, there is shown a cross-sectional view of
 a prior art spar made of 0.073 inches thick 7075-T6 aluminum alloy. This
 spar weighs 0.46 lbs./linear foot. Referring now to FIG. 5, there is shown
 by comparison a cross-sectional view of a titanium-carbon fiber composite
 laminate spar according to the present invention made with a TIMETAL.RTM.
 15-3 shell 62 reinforced with carbon/epoxy strips 64 bonded to the flange
 sections 66 to form a composite laminate structure. This composite
 laminate spar has a strength and stiffness that is similar to the prior
 art spar shown in FIG. 4, but the composite laminate spar weighs only 0.20
 lbs./linear foot, therefore offering a 50% weight reduction.
 Similarly, plates using in aircraft can be constructed from the
 titanium-carbon fiber composite laminates of the present invention where
 the carbon composite layer or layers are in an isotropic pattern or in a
 unidirectional pattern, as suitable for the anticipated stress on the
 plate. For example, referring now to FIG. 6, there is shown a
 cross-sectional view of a five layer titanium-carbon fiber composite
 laminate plate according to the present invention for use in an aircraft.
 The plate 68 is 0.090 inches thick and has unidirectional tensile strength
 of over 300,000 psi. Layers 70, 74 and 78 are beta titanium alloy 0.010
 inches thick. Layers 72 and 76 are unidirectional graphite composite. The
 layers 70 and 78 have one side of their surfaces, 71 and 77 respectively,
 prepared by sandblasting, platinum coating and priming before bonding to
 their respective graphite composite layers 72 and 76, respectively. The
 beta titanium alloy layer 74 has both its surfaces prepared by the above
 method before bonding to the graphite composite layers 72 and 76.
 The yield strength, density and specific strength of the plate 68 are
 compared against prior art materials in FIG. 1. As can be seen in FIG. 1,
 the beta titanium alloy/unidirectional graphite composite has the highest
 specific strength, 3,040,000 inch-lbs. per inch in the L direction, which
 is superior to other materials commonly used for aerospace part
 construction. This high specific strength allows reduction in the overall
 weight by approximately 60% for airplane skin. By comparison, an aluminum
 plate would have to be 41/2 times as thick and 41/2 times as heavy to have
 the same specific strength.
 Although the present invention has been described in considerable detail
 with reference to certain preferred embodiments, other embodiments are
 possible, as will be understood by those with skill in the art with
 reference to this disclosure. Therefore, the spirit and scope of the
 appended claims should not be limited to the description of the preferred
 embodiments.