Form factored compliant metallic transition element for attaching a ceramic element to a metallic element

A ceramic element, e.g., a sapphire dome, is joined to a metallic element, e.g., a vehicle body comprising a titanium alloy, by an attachment structure, e.g., comprising niobium. The attachment structure comprises: (1) a form-factored, compliant metallic transition element having a “C” shape; (2) a first joint material connecting an upper portion of the transition element to the ceramic element; and (3) a second joint material connecting a lower portion of the transition element to the metallic element. A method is provided for attaching the ceramic element to the metallic element, using a single brazing operation. The presence of the attachment structure further minimizes the stresses related to the different coefficients of thermal expansion in the ceramic/attachment/titanium connection.

TECHNICAL FIELD

The present invention relates generally to attaching or securing a ceramic element to a metallic element, and, more particularly, to a vehicle having a ceramic dome and to the attachment or securement of the ceramic dome to the vehicle.

BACKGROUND ART

Outwardly-looking radar, infrared, and/or visible-light sensors built into vehicles such as aircraft or missiles are usually protected by a covering termed a dome or radome. The dome serves as a window that transmits the radiation sensed by the sensor. It also acts as a structural element that protects the sensor and carries aero-dynamic loadings. In many cases, the dome protects a forward-looking sensor, so that the dome must bear large aerostructural loadings.

In one embodiment, an infrared seeker system for missile design generally employs as the dome a protective non-opaque surface to protect its inherently delicate components. Typical applications for this protective surface are semi-spherical or semi-aspherical (conformed) ceramic domes. One popular material for missile applications in the infrared wavelength band is sapphire (a form of Al2O3). These sapphire domes must be located to the missile body by one or more attachment mechanisms.

A common practice for these attachment mechanisms is kinematic mechanical clamps or locating devices combined with high temperature silicon glue. Failure in these joints can occur due to missile flight dynamics, causing thermal and stress conditions exceeding the operational strength of the joint. Over the last few years, Raytheon engineers have devised techniques and processes to replace the silicon joints with brazed sapphire dome assemblies; see, e.g., U.S. Pat. No. 5,758,845, entitled “VEHICLE HAVING A CERAMIC RADOME WITH A COMPLIANT, DISENGAGEABLE ATTACHMENT”, issued on Jun. 2, 1998, to Wayne Sunne et al; U.S. Pat. No. 5,884,864, entitled “VEHICLE HAVING A CERAMIC RADOME AFFIXED THERETO BY A COMPLIANT METALLIC TRANSITION ELEMENT”, issued on Mar. 23, 1999, to Wayne Sunne et al; U.S. Pat. No. 5,941,479, entitled “VEHICLE HAVING A CERAMIC RADOME AFFIXED THERETO BY A COMPLIANT METALLIC “T”-FLEXURE ELEMENT”, issued on Aug. 24, 1999, to Wayne L. Sunne et al; U.S. Pat. No. 6,123,026, entitled “SYSTEM AND METHOD FOR INCREASING THE DURABILITY OF A SAPPHIRE WINDOW IN HIGH STRESS ENVIRONMENTS”, issued on Sep. 26, 2000, to James H. Gottlieb; U.S. Pat. No. 6,241,184, entitled “VEHICLE HAVING A CERAMIC RADOME JOINED THERETO BY AN ACTIVELY BRAZED COMPLIANT METALLIC TRANSITION ELEMENT”, issued on Jun. 5, 2001, to Wayne Sunne et al. The foregoing patents are all assigned to the same assignee as the present application. Brazed sapphire dome assemblies have out-performed earlier state-of-the-art assemblies.

Nevertheless, improvements are continually sought to further reduce stresses related to the different coefficients of thermal expansion in the sapphire (dome)/niobium (transition)/titanium (body) connection.

DISCLOSURE OF INVENTION

In accordance with one embodiment of the present invention, a combination of a ceramic element joint to a metallic element by an attachment structure is provided. The attachment structure comprises:(a) a form-factored, compliant metallic transition element having a “C” shape;(b) a first joint material connecting an upper portion of the transition element to the ceramic element; and(c) a second joint material connecting a lower portion of the transition element to the metallic element.

In accordance with another embodiment of the present invention, a vehicle having a ceramic dome is provided, comprising:(a) a vehicle body having an opening therein;(b) the ceramic dome sized to cover the opening of the vehicle body; and(c) an attachment structure joining the dome to the vehicle body to cover the opening, the attachment structure comprising(i) a form-factored, compliant metallic transition element having a “C” shape,(ii) a first joint material connecting an upper portion of the transition element to the ceramic dome, and(iii) a second joint material connecting a lower portion of the transition element to the vehicle body.

Further in accordance with the present invention, a form-factored, compliant metallic transition element having a “C”-shape is provided for connecting a ceramic element to a metallic element. The transition element combines a transition in coefficient of thermal conductivity and stress relief in one element.

Yet further in accordance with the present invention, a method is provided for securing a ceramic element to a metallic element, each having a different coefficient of thermal expansion, with a transition element to permit flexibility and absorb expansion. The method comprises the steps of:(a) providing as the transition element a form-factored compliant metallic transition element having a “C”-shape;(b) positioning the transition element between the ceramic element and the metallic element;(c) in either order, disposing a first brazing alloy disposed between the ceramic element and the transition element and disposing a second brazing alloy between the transition element and the metallic element, the first and second brazing alloys having substantially the same brazing temperature; and(d) brazing the first and second brazing alloys at the same time to complete securing the ceramic element to the metallic element.

Also in accordance with the present invention, a method is provided for preparing the vehicle having the ceramic dome affixed thereto. The method comprises the steps of:(a) providing the vehicle body having an opening therein;(b) providing the ceramic dome sized to cover the opening of the vehicle body;(c) positioning the compliant metallic aero-shield “C”-shaped flexure element disposed structurally between the dome and the body;(d) affixing the dome to the vehicle body using a first brazing alloy disposed between the dome and the flexure element and a second brazing alloy disposed between the flexure element and the vehicle body, the first and second brazing alloys having substantially the same brazing temperature so that securing the ceramic dome to the vehicle body is accomplished in a single brazing operation; and(e) brazing the first and second brazing alloys at the same time to complete securing the dome to the vehicle body.

The structure disclosed and claimed herein further minimize the stresses related to the different coefficients of thermal expansion in the ceramic sapphire (dome)/niobium (transition)/metallic titanium (body) connection.

BEST MODES FOR CARRYING OUT THE INVENTION

FIG. 1depicts a vehicle, here illustrated as a missile20, having a dome or radome21attached thereto. The dome21is forwardly facing as the missile flies and is therefore provided with a generally ogival shape that achieves a compromise between good aerodynamic properties and good radiation transmission properties. The missile20has a missile body22with a forward end24, rearward end26, and a body axis27. The missile body22is generally cylindrical, but it need not be perfectly so. Movable control fins28and an engine30(a rearward portion of which is visible inFIG. 1) are supported on the missile body22. Inside the body of the missile are additional components that are not visible inFIG. 1, are well-known in the art, and whose detailed construction are not pertinent to the present invention, including, for example, a seeker having a sensor, a guidance controller, motors for moving the control fins, a warhead, and a supply of fuel.

Infrared Seeker Technology for missile designs generally employs a protective non-opaque surface to protect its inherently delicate components. Typical applications for this protective surface are semi-spherical or semi-aspherical (conformed) ceramic domes. One popular material for missile applications in the infrared wavelength band is sapphire. These sapphire domes must be located to the missile body by one or more attachment mechanisms. Common practice for these mechanisms is kinematic mechanical clamps or locating devices combined with high temperature silicon glue. Failure in these joints can occur due to missile flight dynamics, causing thermal and stress conditions exceeding the operational strength of the joint. Over the last few years, Raytheon engineers have devised techniques and processes to replace the silicon joints brazed sapphire dome assemblies. These assemblies have out-performed the previous state of the art.

An example of a state-of-the-art brazed sapphire dome assembly design is depicted in FIG.2. In that design, a sapphire dome is brazed to a niobium washer and is in turn brazed to a titanium flexure. The brazed assembly is then protected by an aero-shield in order to protect the double brazed joint from the aero-thermal environment inherent in missile flight. An air gap insulates the inner surface exposed to the sensor from the outer surface exposed to the air stream.

FIG. 2illustrates a region at the forward end24of the missile body22, where the dome21attaches to the missile body22. The dome21has an inside surface32, an outside surface34, and a lower margin surface36extending between the inner surface32and the outer surface34. The lower margin surface36is generally perpendicular to the body axis27. The dome21is made of a ceramic material, typically, sapphire, a form of aluminum oxide. For structural reasons, the dome21is typically fabricated with a crystallographic c-axis38of the sapphire generally (but not necessarily exactly) perpendicular to the margin surface36. Thus, in the region of the dome21near to the margin surface36, the crystallographic a-axis40of the sapphire is generally (but not necessarily exactly) perpendicular to the inner surface32and to the outer surface34. However, for some applications, the crystallographic orientation of the sapphire may be other than along the a- or c-axis, in order to provide certain structural advantages for aerodynamic loading, such as disclosed, for example, in U.S. Pat. No. 6,123,026, issued Sep. 26, 2000.

The most forward end of the missile body22defines a nose opening42, which in this case is substantially circular because the missile body is generally cylindrical. An attachment structure44joins the dome21to the missile body22in order to cover and enclose the opening42. The attachment structure includes a compliant “T”-flexure element46, which is an integral part of the missile body22. The “T”-flexure element46has the form of a ring that extends around the entire opening42, but is shown in section in FIG.2.

In section, the “T”-flexure element46has a substantially T-shape, and comprises an elongated compliant arm region48that extends generally parallel to the body axis27of the missile20. The arm region48is secured at one end48ato the missile body22and, in fact, is integral with the missile body. A crossbar region50, secured to the opposite end48b, is perpendicular to the arm region48and thence generally perpendicular to the body axis27. The arm region48and the crossbar region50are integrally formed as part of the missile body22. The arm region48and the crossbar region50preferably extend completely around the circumference of the ring of the “T”-flexure element46. Essentially, the missile body22is thinned in the area of the arm region48so as to provide flexure, as described more fully below. The thinning of the arm region48is conventional and forms no part of the present invention.

The dome21is joined to the “T”-flexure element46at a first attachment, through a niobium-containing washer47. The first attachment is preferably a first brazed butt joint54between an upper surface47aof the niobium washer47of the “T”-flexure element46and the lower margin surface36of the ceramic dome21. The first brazed butt joint54is preferably formed using an active brazing alloy that chemically reacts with the material of the dome21during the brazing operation.

In forming this butt joint54, care is taken that the brazing alloy contacts only the lower margin surface36of the dome21, and not its inside surface32or its outside surface34. The molten form of the active brazing alloy used to form the butt joint54can damage the inside surface32and the outside surface34of the dome, which lie perpendicular to the crystallographic a-axis40of the sapphire material. The lower margin surface36, which lies perpendicular to the crystallographic c-axis38of the sapphire material, is much more resistant to damage by the active brazing alloy. The use of the butt joint only to the lower margin surface36of the sapphire dome thus minimizes damage to the sapphire material induced by the attachment approach.

The use of a butt joint to join the dome21to the “T”-flexure element46is to be contrasted with the more common approach for forming joints of two structures, a lap or shear joint. In this case, the lap joint would be undesirable for two reasons. The first, as discussed in the preceding paragraph, is that the lap joint would necessarily cause contact of the brazing alloy to the inside and/or outside surfaces of the dome, which are more sensitive to damage by the molten brazing alloy. The second is that the lap or shear joint would extend a distance upwardly along the inside or outside surface of the dome, reducing the side-viewing angle for the sensor that is located with the dome. That is, the further the opaque lap joint would extend along the surface of the dome, the less viewing angle would be available for the sensor. In some applications, this reduction of the side-viewing angle would be critical.

The niobium-containing washer47is joined to the “T”-flexure element46at a second attachment. The second attachment includes a second brazed butt joint58between a lower surface47bof the washer47and an upper surface50aof the crossbar region50.

The missile body22is preferably made of a metal such as a titanium alloy. The titanium alloy of the missile body22and the sapphire of the dome21have different coefficients of thermal expansion (CTE). When the missile20is heated and cooled during fabrication or service, this difference in thermal expansion coefficients causes the total expansion of the dome21and the missile body22to be different. This difference would ordinarily produce thermally induced stresses in the dome21and the missile body22. The thermally induced stresses have relatively small effects on the metallic missile body structure, but they can produce significant damage and reduction in failure stress in the ceramic material of the dome21. The present approach of the combination of the “T”-flexure element46and niobium-containing washer47avoids or minimizes such thermally induced stresses.

The “T”-flexure element46is made of the same metal or metal alloy as the missile body22. The arm region48is made relatively thin, so that it can bend and flex to accommodate differences in the coefficients of thermal expansion of the missile body22and the dome21. Stated alternatively, the thermally induced stresses are introduced into the arm region48of the “T”-flexure element46and not into the dome21. Further, the niobium-containing washer47acts as a CTE mismatch bridge between the sapphire dome21and the titanium body22.

An aero ring60is brazed to the missile body22with a braze joint62and is used to protect the “T”-flexure element46and the niobium-containing washer47against aerodynamic stresses and temperatures during flight. The aero ring60may be spaced from the niobium-containing washer47, as shown inFIG. 2, or may be butted against a portion of the bottom surface of the washer and sealed with a heat-resistant polymer, such as polysulfide (not shown).

In accordance with the present invention, a sapphire dome is secured to a titanium body using a form-factored niobium transition flexure.FIG. 3shows the sapphire dome21, preferably brazed to the form-factored niobium transition flexure160, using a first braze alloy. The dome21may be a conformal optical dome or a non-conformal optical dome. The teachings herein are not limited to the type of dome employed in the missile20.

The form-factored niobium transition flexure160is preferably brazed to the titanium body, here, dome mount22, using a second braze alloy. Incusil ABA braze alloy is used as the first braze alloy, while Incusil-15 is used as the second braze alloy. Incusil ABA and Insusil-15 are registered tradenames of WESGO Inc. Incusil ABA is an active braze alloy having a composition, in weight percent, of about 27.25 percent copper, about 12.5 percent indium, about 1.25 percent titanium, and the balance about 59 wt % silver, while Incusil-15, also an active braze alloy, has a composition, in weight percent of 61.5 percent silver, 23.5 percent copper, and 15 percent indium.

Whereas the specific braze alloys listed above have been optimized for this particular application, any brazing material within the active silver braze alloy family may be used for the first braze and any brazing material within the titanium doped active silver alloy family may be used for the second braze. The physical performance requirements of the assembly drive optimization to a particular alloy within the respective family of alloys for the brazed joints.

The design of the present invention employs the conventional current state-of-the-art features: thin niobium washers154and162as the transition elements and the separate aero-shield160(form-factored niobium transition element). A titanium heat shield170serves as a heat baffle, due to the extreme aero-thermal environment. The titanium heat shield170is incorporated as a feature of the dome mount (or missile body)22in order to simulate the air space42formed by the prior art titanium aero-shield60, braze joint62, and titanium flexure48. This air space is required in order to create an air pocket insulation between the high operational temperatures of the outer missile body22and the intrinsically delicate electronic parts, including the seeker, within the missile body22and the dome21.

The niobium transition element47ofFIG. 2is essentially stretched into the “C”-shaped transition element160of FIG.3. This reconfiguration of the niobium washer47into the “C”-shaped washer160allows it to perform the same functions as both the flexure48and aero-shield60ofFIG. 2, along with its original purpose of providing a stress-absorbing transition element47between the titanium dome mount22and the dome21.

The major change from the state-of-the-art design to the design of the present invention exists in the aero-shield160replacing the aero ring60, thereby obviating the additional second braze location58(in FIG.2). In the present design, the niobium aero-shield160is used to secure the dome21to the missile body22.

The shape of the niobium aero-shield160is contoured to match the shape of the vehicle, here, missile20, thereby eliminating the need for a secondary missile shield (element60in FIG.2). The niobium aero-shield160may be formed by a number of different methods, including, but not limited to, spin-forming, machining, or die-forming. By “form-factored” is meant that the niobium aero-shield160is preformed in a purpose-efficient shape. As used in a missile20, this means that the aero-shield160is formed in a shape that is useful as part of the missile design.

The niobium aero-shield160is formed as a “C”-channel. Preferably, the aero-shield160has a generally flat upper connector portion160bhaving an inner annulus and an outer annulus, a generally flat lower connector portion160chaving an inner annulus and an outer annulus, and a flexure portion160aconnecting the upper portion and the lower portion at the outer annulus of each.

The flexure portion160ahas a relatively thin cross-section in the flexure region160a, from about 0.010 to 0.025 inch (0.254 to 0.635 mm), preferably about 0.015 inch (0.381 inch). The top connector portion160bis somewhat thicker, but still relatively thin, in order to reduce stress on the dome21. The thickness of the top connector portion160branges from about 0.020 to 0.030 inch (0.508 to 0.762 mm). The bottom connector portion160cis somewhat thicker still, and ranges from about 0.035 to 0.045 inch (0.889 to 1.143 mm).

The niobium aero-shield160is self-locating. That is to say, the gap between the niobium aero-shield160and the titanium turret22ahas been designed to be self-locating. Because the coefficient of expansion for Ti is greater than that for Nb, the gap has been designed so that at the braze temperature, the fit is at or close to line-to-line diametrically. This causes the inside diameter of the Nb aero-shield160(initially larger, but slower growing) to be forced concentric with the outside (initially smaller, but faster growing) diameter of the Ti turret22a. Thereby, the thermal cycle of the braze operation centers the Nb aero-shield160on the Ti turret22a.

The braze alloy disks used to form the braze joints154,162are prefabricated rings of the appropriate annular diameter and are about 0.002 inch (0.051 mm) thick.

It is known that dissimilar materials possess dissimilar growth rates under thermal load. The rate of growth for particular materials is represented by its coefficient of thermal expansion (CTE). If two dissimilar metals are welded/brazed/-glued together and subsequently thermally cycled, a sheer stress directly related to the difference in material CTE will result. Sapphire and niobium have very similar CTEs. This results in the sapphire and niobium growing at very similar rates during the thermal changes, occurring during both flight and the braze process. Therefore, the brazed joint between the sapphire dome21and the niobium aero-shield160sees little sheer stress under heat cycling.

Titanium has a significantly higher CTE than the sapphire and the niobium. To reduce the stress at the titanium/niobium joint, the design employs a flexure allowing a prescribed displacement to reduce the stiffness of the joint. As the assembly is heat cycled, the titanium begins to out-grow the sapphire and niobium. Consequently, the thin flexure160abegins to displace, thereby reducing and controlling the stress at the niobium/titanium joint.

To fabricate the missile20having the dome21joined to the missile body22, the missile body22is provided, together with (1) the heat shield170, (2) the “C”-shaped aero-shield/flexure transition element160, and (3) the ceramic dome21. The portion of the missile body22that forms the heat shield170and the turret mount22ais preferably an integral unit as shown in FIG.3and comprises a titanium alloy such as Ti-6A1-4V, having a composition, in weight percent, of 6 percent aluminum, 4 percent vanadium, balance titanium. The aero-shield160is preferably a niobium-based alloy having a composition, in weight percent, of 1 percent zirconium, balance niobium. Other metals or alloys may be employed in place of the niobium-based alloy disclosed, so long as they have a coefficient of thermal expansion that is within about 0.5% that of sapphire and meet other required mechanical properties, such as strength. While examples of such other metals and alloys include tantalum, tantalum-tungsten, and Kovar, such metals and alloys are less preferred than the niobium-based alloy disclosed herein, mainly due to their cost. The niobium-based alloy is further preferred because it is readily available, is easily spin-formed or machined or die-formed, and has a coefficient of thermal expansion relatively close to that of the preferred dome material, sapphire.

The braze alloys154,162described above are relatively low-temperature (approximately 1300° F., or 704° C.) for brazing the aero-shield160to both the ceramic dome21and the turret mount22aof the missile body22. The braze alloys are compatible with the materials of the missile body22and the dome21.

The braze alloys are provided in the form of braze alloy disks, one of which is placed between the aero-shield160(upper connector portion160b) and the ceramic dome21(for forming braze joint154), and the other of which is placed between the aero-shield160(lower connector portion160c) and the turret mount22a(for forming braze joint162). The brazing is accomplished by heating the missile body22, the aero-shield160, and the dome21with the braze alloy washers therebetween, to a brazing temperature sufficient to melt the braze alloy and cause it to flow freely, about 1330° F. (721° C.). The brazing is accomplished in a vacuum of about 8×10−5Torr or less and with a temperature cycle involving a ramping up from room temperature to the brazing temperature of about 1300° F. (704° C.), a hold at the brazing temperature for 9 minutes, and a ramping down to ambient temperature, the total cycle time being about 5 hours.

As noted previously, it is highly desirable that the braze alloy forming the braze joint154not contact the inside surface32or the outside surface34of the dome21, and that the braze alloy only contact the margin surface36. To achieve this end, the first braze alloy154is provided in the form of a flat disk that fits between the margin surface36and the upper connecting surface160b. The volume of the braze element washer is chosen so that, upon melting, the braze material154just fills the region between the margin surface36and the upper connecting surface160b. There is no excess braze alloy to flow onto the surfaces32and34.

Likewise, the second braze alloy forming the second braze alloy joint162is also provided in the form of a flat disk that fits between the lower connecting surface160cand the surface of the turret mount22a.

During the braze operation of joining the ceramic dome21to the missile body22, the aero-shield160is disposed circumferentially around the titanium heat shield170.

The joints154and162are both preferably braze joints, as illustrated. The braze joints are preferred because they form a hermetic seal for the aero-shield160. The hermetic seal prevents atmospheric contaminants from penetrating into the interior of the missile body during storage. It also prevents gasses and particulate material from penetrating into the interior of the missile body during service. Other operable joint structures and joining techniques may be used.

The advantages of the present design over the prior art designs include at least the following:(1) The use of the niobium-based integral flexure and aero-shield160reduces the part count and allows a niobium element to perform three functions: (a) transition element; (b) flexure; and (c) aero-shield.(2) The integration of the niobium transition element into the flexure160results in a lower inherent stress to the dome over niobium washer designs.

The foregoing description has been presented in terms of attachment of a ceramic dome, comprising sapphire, to a metallic body, e.g., a titanium alloy of a missile. However, it will be appreciated by those skilled in this art that the teachings herein are suitably employed for securing other ceramic materials, including alumina, doped alumina (doped with at least one transition metal ion), and other oxides, whether crystalline or non-crystalline, to other metals. In any case, suitable braze materials are used between the transition element and the ceramic element on one side and the metallic element on the other side. The larger the coefficient of thermal expansion between the ceramic material and the metal, then an increase in the length of the flexure element is required in order to permit flexibility and to absorb expansion. However, the determination of the appropriate braze materials and the length of the flexure element are considered to be readily within the ability of one skilled in this art, not requiring undue experimentation, based on the teachings herein.