Abstract:
A redundant ice management system and method to de-ice and anti-ice an aircraft member ( 150 ) is provided. The system comprising a primary ice management sub-system ( 154 ) for providing thermal ice management to the aircraft member and a secondary ice management sub-system ( 172 ) for providing back-up thermal ice management to the aircraft member ( 150 ) in the event of a failure by the primary ice management sub-system ( 154 ). The system provides primary and secondary de-ice and anti-ice capabilities to the aircraft member ( 150 ) before and during airborne operation.

Description:
FIELD OF THE INVENTION 
     The present invention is related to electrical heating systems for the prevention or removal of ice accumulation on the surface of aircraft structural members and, more particularly, to a redundant ice management system for aircraft. 
     BACKGROUND OF THE INVENTION 
     The accumulation of ice on aircraft proprotors, wings and other structural members in flight is a well known danger during low temperature conditions. As used herein, the terms “aircraft members” or “structural members” are intended to refer to any aircraft surface susceptible to icing during flight, including proprotors, wings, stabilizers, engine inlets and the like. Attempts have been made since the earliest days of flight to overcome the problem of ice accumulation. While a variety of techniques have been proposed for removing ice from aircraft before or during flight, many prior systems or techniques experience various drawbacks or possess certain limitations. 
     One approach to ice management that has been used is so-called thermal de-icing. In thermal de-icing, the leading edges, that is, the portions of the aircraft that meet and break the airstream impinging on the aircraft, are heated to prevent the formation of ice or to loosen accumulated ice. The loosened ice is then blown from the structural members by the airstream passing over the aircraft. 
     In one form of thermal de-icing, heating is accomplished by placing an electrothermal pad, including heating elements, over the leading edges of the aircraft, or by incorporating the heating elements into the structural members of the aircraft. Electrical energy for each heating element is typically derived from a generating source driven by one or more of the aircraft engines or transmissions. The electrical energy is intermittently or continuously supplied to provide heat sufficient to prevent the formation of ice or to loosen accumulating ice. 
     With some commonly employed thermal de-icers, the heating elements are configured as ribbons, e.g. interconnected conductive segments, that are mounted on a flexible backing. The conductive segments are separated from each other by gaps, e.g. intersegmental gaps, and each ribbon is electrically energized by a pair of contact strips. When applied to a wing or other airfoil surface, the segments are arranged in strips or zones extending spanwise or chordwise of the aircraft wing, rotor or airfoil. One of these strips, known as a spanwise parting strip, is disposed along a spanwise axis which commonly coincides with a stagnation line that develops during flight in which icing is encountered. Other strips, known as chordwise parting strips, are disposed at the ends of the spanwise parting strip and are aligned along chordwise axes. Other zones, known as spanwise shedding zones, are typically positioned above and below the spanwise parting strip at a location intermediate the chordwise parting strips. Between adjacent zones, a gap, known as an interheater gap, sometimes exists. 
     One known method for de-icing causes electrical current to be transmitted continuously through parting strips so that the strips are heated continuously to a temperature above 32° F. In the spanwise shedding zones, on the other hand, current is transmitted intermittently so that the spanwise shedding zones are heated intermittently to a temperature above about 32° F. 
     While this technique of heating the various zones generally is effective to melt ice (or prevent its formation) without the consumption of excessive current, a problem exists in that melting of ice in the inter-segmental and interheater gaps can be difficult or impossible. Moreover melting of ice on or around the contact strips can also be difficult or impossible. Accumulation of ice in the gaps and on the contact stripe is particularly undesirable because the unmelted ice can serve as “anchors” for ice that would be melted but for the ice accumulated in the gaps or on the contact strips. 
     Another problem with prior thermal-based systems is their lack of reliability. Aircraft members, such as rotors of a helicopter or proprotors of tiltrotor aircraft, undergo much strain and stress associated with aircraft operation. Ongoing use of aircraft inevitably results in some damage to aircraft components. With respect to heating elements integrated within an aircraft member, breaks in blanket circuitry can cause thermal de-icing systems to fail, posing serious risk to aircraft crew and equipment during cold weather operations. And yet another concern with heating element circuitry is the potential for inconsistency, e.g. hot spot or cold spot generation, and larger than acceptable power consumption. 
     Problems may also be encountered where strips are run along the entire length of the aircraft. The size of the ice being shed by the aircraft member can cause a hazard to the aircraft&#39;s fuselage. If the particle of ice is too large, it could hit and may even penetrate the fuselage. 
     SUMMARY OF THE INVENTION 
     In response to the foregoing concerns, the present invention provides a new and improved thermal ice management system for aircraft structural members. Specifically, the present invention provides a secondary section having secondary anti-ice elements and secondary de-ice zones which provide thermal ice management to aircraft structural members. 
     The redundant ice management system of the present invention, includes a primary ice management sub-system that provides thermal ice management to aircraft structural members and a secondary ice management sub-system that provides back-up thermal ice management to aircraft structural members in the event of a failure by the primary ice management sub-system. 
     Further novel aspects of the present invention are found with the incorporation and use of separate zones within the primary and secondary sub-systems, integration of the redundant ice management systems with a controller and the integration of the controller with atmospheric, structural and system monitoring capabilities. 
     The present invention also provides a method for managing the formation of ice on aircraft structural members with an ice management system having primary and secondary ice management sub-systems that includes monitoring aircraft structural members and atmospheric conditions for ice formation on the aircraft&#39;s structural members, activating primary ice management systems in response to an indication of ice formation on the aircraft&#39;s structural members, monitoring the primary ice management systems to determine its operational readiness and efficiency and activating the secondary ice management system in response to monitoring of the primary ice management if the primary ice management system fails operational readiness and efficiency requirements. 
     One advantage of the present invention is that it provides for a backup ice management scheme in the event of a failure by the primary system. By providing primary and secondary sub-system elements, heat is effectively and efficiently generated throughout the aircraft member regardless of primary system failure. Sections of the primary and secondary sub-system elements are oriented spanwise and chordwise along the aircraft&#39;s structural member in a manner that can provide adequate surface coverage for thermal management operations. 
     Another advantage of the present invention is that it optimizes element dimensions, such that primary and secondary sub-system sections promote efficient heating along the entire targeted area and minimizes the amount of overlapping that is required to gain desired heat distribution for thermal ice management. 
     Yet another advantage of the present invention is that it eliminates cold spots which can arise on and around aircraft structural member through selective activation of heating elements disposed along a structural member. 
     Another advantage of the present invention is that it affords highly desirable levels of heating while using a minimum amount of power. More specifically, by sequentially heating spanwise shedding areas, power consumption is minimized by the controller without sacrificing de-icing capabilities. Additionally, flexible control of the primary and secondary elements maximizes de-icing capability. In particular, as flight conditions change, the interval during which each systems elements are heated can be varied by an onboard controller. 
     Another advantage of the invention is the stepwise employment of eight chordwise zones of de-icing in the spanwise direction from the rotor blade tip to root, rather than full span chordwise zones on the upper and lower rotor surfaces. Resulting ice pieces are therefore smaller and don&#39;t pose as great a risk in penetrating the aircraft&#39;s fuselage. 
    
    
     BRIEF DESCRIPTION OF THE DRAWINGS 
     For a more complete understanding of the present invention, including its features and advantages, reference is now made to the detailed description of the invention, taken in conjunction with the accompanying drawings of which: 
     FIG. 1 is a partial perspective view of a prior art air foil having a thermal de-icer mounted along the airfoil&#39;s leading edge; 
     FIG. 2 is a top plan view of a prior art thermal de-icer; 
     FIG. 3 is a partial, broken-away top plan view of a prior art thermal de-icer mounted on a structural member; 
     FIG. 4 is a vertical cross-sectional view of the layout for a prior art thermal de-icer taken along the stagnation line of FIG. 3; 
     FIG. 5 is cross-sectional view of a rotor blade of a helicopter or a proprotor of a tiltrotor aircraft present invention can be utilized; 
     FIGS. 6A-6B are top plan views of a schematic layout for the primary and secondary heating systems of the present invention; and 
     FIG. 7 is a block diagram view of the system components for the present invention. 
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     While the making and using of various embodiments of the present invention is discussed in detail below, it should be appreciated that the present invention provides many applicable inventive concepts which can be embodied in a wide variety of specific contexts. The specific embodiments discussed herein are merely illustrative of specific ways to make and use the invention, and do not delimit the scope of the invention. 
     The present invention is directed toward thermal control over the development of ice on aircraft structural members such as proprotors and wings. The invention involves incorporation of heater blanket technology as used in aircraft to remove ice from the leading edge of the aircrafts blade or proprotor. The blanket technology of the present invention includes separately controlled sub-systems, referred to as the primary heating system and the secondary heating system throughout this description. The purpose for having redundant systems is to provide a backup system for the aircraft and its crew if the primary system&#39;s heater elements fail. A secondary system allows continued operations with secondary de-ice or anti-ice management of an aircrafts blades and rotors. 
     Aircraft having thermal ice removal systems, may include an anti-ice zone that is heated so that ice is never allowed to forms and a de-ice zone wherein ice is allowed to form to a certain thickness and then is removed when heater elements are activated, bringing the surface temperature, through an abrasion strip, up to a point where the surface tension is reduced and the ice will fall away, be blown away by air flow over the aerodynamic surface or by the centrifugal force caused by rotor rotation. 
     Referring to FIG. 1, a thermal de-icer  10  according to one implementation by the prior art is shown mounted on a structural member  11  in the form of a wing. As is known, the structural member  11  includes a chordwise axis and a slantwise axis. During flight, airflow impinges a leading edge  13  of the structural member  11 , and a number of stagnation points develop, forming a stagnation line or axis, which stagnation line varies during flight conditions. 
     The de-icer  10  is mounted symmetrically about the stagnation line which would be most commonly encountered during icing conditions. Due to the sweep of the structural member  11  upon which the de-icer  10  is employed, a pair of chordwise disposed or side edges of the de-icer  10  have a chevron shape when the de-icer  10  is flat. As will be appreciated by those skilled in the art, configuring the side edges in this manner allows for two of de-icers  10  to be placed side-by-side, along the leading edge  13 , without forming a gap between the two de-icers  10 . For a structural member  11  with no sweep, the side edges would be perpendicular with the stagnation line when the de-icer  10  is flat. In the following discussion, the operation of a single de-icer  10  will be discussed. It should be recognized, nonetheless, that commonly a number of de-icers  10  would be mounted adjacent to one another along the leading edge  13  of the structural member  11 . 
     FIG. 2 illustrates in further detail the prior art thermal de-icer  10  which includes a plurality of elements or ribbons  12 . The elements  12  are typically mounted on a flexible backing  15 . Then elements are arranged to provide a stepwise parting strip  14 , chordwise parting strips  16 , and stepwise shedding zones  18 . Current is transmitted to the elements  12  by way of contacts  20 - 23 . Contacts  20 - 23  include four pairs of contact pads, four of which pads are disposed on one end of the de-icer  10  and the other four of which are disposed on an opposing end of the de-icer  10 . In operation, voltage differences are established between the pad pairs so that current flows through each of the elements  12 . 
     Interheater gaps  24  are disposed between the various zones  14 ,  16  and  18 . The elements  12  are defined by interconnected conductive segments  26 , which conductive segments  26  are aligned along axes that are parallel with either the stagnation line or chordwise axes of the structural member  11 . Each pair of conductive segments  26  is interconnected by a turn  28  and defines an inter-segmental gap  30 . 
     In operation, current is transmitted continuously to the spanwise and chordwise parting strips  14 ,  16  so that heat is generated continuously therein. Heat is generated continuously in the spanwise parting strip  14  since ice that accumulates adjacent to the stagnation line, such as rime ice, tends to be most difficult to melt. Current is transmitted intermittently to the spanwise shedding zones  18  so that heat is generated intermittently therein. 
     One object of the de-icer  10  is to melt all of the ice that accumulates adjacent to the elements  12 , but in practice certain problems arise. First, with heating or de-icing systems such as de-icer  10 , ice can accumulate in the interheater gaps  24  as well as in the inter-segmental gaps  30 . More specifically, during operation, very little current flows in the outer portions or corners of the turns  28  so that even when, for example, the turns  28  of one of the elements  12  are positioned close to the turns  28  of another of the elements  12 , there still is no practical way to transfer heat from the one set of turns  28  to the other set of turns. Second, in common prior art arrangements of de-icer  10 , no heat is supplied to contacts  20 - 23 . In particular, the contact pads of contacts  20 - 23  are much wider than typical conductive segments  26  and are attached to a heavy lead wire having a relatively large cross-sectional area. Thus, the contact pads dissipate relatively little energy and become cold spots, upon which ice accumulates. Moreover, the contact pads serve as “anchors” for ice which would have melted but for the cold spots generated by the contacts  20 - 23 . Third, the interheater gaps  24  between the chordwise parting strips  16  and the spanwise shedding zones  18  are particularly difficult to heat. More specifically, the outside corners of the turns  28  disposed near the chordwise parting strip  16  are angled to accommodate for chevron-shaped edges of the de-icer  10 . 
     In operation, current does not flow efficiently in these angled corners and the resulting cold spot(s) can make the task of sufficiently heating the interheater gaps  24  even more difficult. Finally, some of the conductive segments  26  are too short in length to provide adequate heating. It has been found that when the conductive segments  26  are too short, current flux density is such that an undesirable heating pattern is achieved in the element  12 . 
     It is believed that de-icer  10 , while certainly more efficient than many known thermal de-icers, is incapable of minimizing cold spots. That is, even if cold spots could be eliminated in the interheater gaps  24  by generating more heat in the elements  12 , the de-icer  10  still would consume undesirably high levels of power. Moreover, generation of more heat would not necessarily allow for melting in the region of the contacts  20 - 23  or in certain of the turns  28  formed near the chordwise parting strips  16 . 
     Referring to FIGS. 3 and 4, a partial plan view and perspective view, respectively, of a prior art thermal de-icing system is shown. The de-icer  40  provides heat to the interheater gaps  24  and the inter-segmental gaps  30  as well as to the contacts  20 - 23  (as shown in FIG.  2 ). The de-icer  40  is mounted along the leading edge  13  (FIG. 1) of the structural member  11 . The structural member  11  is typically a composite material, but, in other examples, could be a metal, such as aluminum. Referring to FIG. 4, the de-icer system  40  may includes spanwise parting strips  44 , chordwise parting strips  45  and spanwise shedding zones  46 , each mounted on a flexible backing (not shown). The spanwise parting strip  44  preferably is mounted along an axis which is coincidental with a stagnation line most commonly encountered during icing conditions. The strips  44 ,  45  and the zones  46  include conductive elements or ribbons  50  which are positioned along either a spanwise or a chordwise axis. The elements  50  preferably are configured in serpentine patterns. 
     Referring to FIG. 3, current is transmitted to the elements  50  by way of contacts  51 , which contacts  51  are connected to the elements  50 . Contacts  51  include pairs of contact strips or pads, each of which strip is connected to an end of element  50  and includes a substantial portion disposed remotely of strips  44 ,  45  and zones  46 . Only one contract strip is shown for each of the elements  50  in FIG.  3 . It should be appreciated that such overlap eliminates cold spots which can exist in interheater gaps  50  during the heating of elements  50 , and facilitates more desirable heat distribution between elements  50 . 
     Cold spots, which can function as ice anchors, commonly form in the area covered by the contacts  51 . Referring again to FIG. 3, local cold spots attributable to the contacts  51  are eliminated by overlapping the contacts  51  with the chordwise parting strip  45 . Under one alternative technique for eliminating cold spots attributable to the contacts  51 , the contacts  51  are folded under the elements  50  subsequent to mounting and etching of the elements  50  and contacts  51  on either of backings  47 ,  48 . Under another alternative for eliminating cold spots, the contacts  51  are overlapped by a spanwise parting strip  44  or a spanwise shedding zone  46 . 
     When the de-icer  40  is attached to an upper surface of structural member  11 , lead wires are coupled to contacts  51 . During installation lead wires are extended from the electrical system of the aircraft and through the leading edge  13  to the contact means  51 . It also can be appreciated that chordwise parting strips  45  have contacts (not shown) which in one embodiment can be disposed under portions of the one or more spanwise parting strips  44 . 
     Referring to FIGS. 5A-5B, therein is depicted cross-sectional views of a proprotor  100  as representative of one type of aircraft member utilizing the present invention wherein primary and secondary heating systems would be incorporated into its leading edge  102 . 
     Proprotor blade  100  is constructed from a plurality of fiberglass skins such as fiberglass skin  104  and fiberglass skin  106  which form the aft body  108  shape of the blade  100 . Surrounding the leading edge  102 , blade  100  is covered by an abrasion strip assembly  110  that may be titanium or other suitable material and the heater blanket  112  which is bonded with adhesive to the blade spar  114 . In addition, on the abrasion strip assembly  110 , a nose cap  116  is positioned at the outermost edge of the leading edge  102 . Disposed within rotor blade  100  at the leading edge of the spar  114  is an inertia weight  118 . 
     As best seen in FIG. 5B, the abrasion strip assembly  110  is made up of the abrasion strip  120 , nose cap  116  and the heater blanket  112 . The heater blanket  112  is disposed between fiberglass layers  122 ,  124  and includes a fiberglass layer  126  therein. Disposed between the fiberglass layers  124  and the fiberglass layer  126  is the primary heating system  128 . Disposed between the fiberglass layer  126  and the fiberglass layer  122  is the secondary heating system  130 . The primary heating system  128  includes an anti-ice zone  132 . The primary heating system  128  also includes a plurality of de-ice zones, such as a de-ice zone  134  positioning aft of the anti-ice zone  132  and on the upper surface of the abrasion strip assembly  110  and a de-ice zone  136  aft of the anti-ice zone  132  and on the lower surface of the abrasion strip assembly  110 . Likewise, secondary heating system  130  includes an anti-ice zone  138  at the leading edge  102  of abrasion strip assembly  110  and a plurality of de-ice zones such as the de-ice zone  140  and the de-ice zone  142 . 
     Referring to FIGS. 6A and 6B, therein are depicted spanwise schematic layouts of various layers of the proprotor proximate the leading edge. In FIG. 6A, proprotor section  150  has been unfolded about axis  152  which represents the leading edge of the proprotor such that the illustrated layer containing the primary heating system  154 . The primary heating system  154  is divided into eight de-ice zones, specifically zones  156 - 170 , starting at the tip of the proprotor section  150  and being of substantially equally-sized. The zones  156 - 170  cover the leading edge of the proprotor spanwise towards the inboard section of the proprotor  150 . The secondary heating system  172  is depicted in FIG.  6 B and has substantially overlapping coverage with the primary heating system  154 . The secondary heating system  172  is divided into four generally equally spaced zones  174 - 180 . It should be appreciated that neither the primary heating system  154  nor the secondary heating system  172  are restricted to the number of zones that may be implemented on an aircraft member. 
     Both the primary heating system  154  and the secondary heating system  172  have anti-ice zones  182 ,  184 , respectively that are incorporated into a portion of the center of the leading edge of the proprotor section  150 . The anti-ice zones are preferably incorporated from about half the span of the proprotor&#39;s span to the tip of the proprotor  150 , and are less than an inch wide. The primary anti-ice zone  182  is on the very leading edge of the proprotor section  150 . Underneath the primary anti-ice zone  182  is the secondary anti-ice zone  184 , as best seen in FIG.  5 . 
     The circuits for the primary heating system  154  and secondary heating systems  122  are completely separate. The primary anti-ice and de-ice systems share a common bus  190 . The secondary anti-ice and de-ice systems share a common bus  192 . The primary de-ice zones  156 - 170  are each provided electrical current via primary de-ice contacts and buses  194 . The secondary de-ice zones  174 - 180  are each provided electrical current through their respective de-ice contacts and buses  196 . The primary anti-ice zone is provided electrical current through contact/bus  198 , and the secondary anti-ice zone is provided electrical current through contact/bus arrangement  200 . A 3-phase power system is preferably used by the primary and secondary heating systems  154 ,  172 . In order to provide absolute system redundancy, it is desirable to have separate primary and secondary power sources for the separate circuitry. 
     Referring to FIG. 7, a programable controller  210  manages the entire system. Power to the zones  156 - 170  of the primary heating system  154  and the zones  174 - 180  secondary heating system  172  are cycled by the controller  210 . The secondary heating system  172  is invoked by the controller  210  when failure of the primary heating system  154  is sensed by sensors  212 . The detection sensors  212  inform the controller  210  of a malfunction, a short, an open, or a change in the resistance of significant amount and it will shut that particular zone down. The controller  210  may cycle all the other primary de-ice zones. Alternatively, the controller  210  may completely by-pass the primary heating system  154  and invoke the full power of the secondary heating system  172 . 
     A dedicated system controller  210  is best-suited for monitoring of sensors  212  and circuit management operations for the primary heating system  154  and secondary heating system  172 . A dedicated controller  210  senses a problem, e.g., a short circuit in one of the zones, and can bypass the problem. The controller&#39;s zone cycling may be sophisticated depending on its programming. Sensing may also take into account, for example, depending on the severity of the ice condition, temperature and size of droplets (e.g., temperature, droplet size, number of droplets, formation of ice, speed of ice formation). The controller  210  will manage the duration that a particular zone is on based on the monitored conditions. Typically, a de-ice zone is not heated for more than 15 seconds. The controller  210  can be programmed to automatically manage the power systems for the aircraft. The controller  210  can be responsible for power conservation. Under normal circumstances, the secondary heating system  172  would only operate after failure of the primary heating system  154 . The pilot, however, may be provided the option to override the heating controller functions as indicated at  214 . 
     In the foregoing description, it will be readily appreciated by those skilled in the art that modifications may be made to the invention without departing from the concepts disclosed herein. Such modifications are to be considered as included in the following claims unless those claims, by their language, expressly state otherwise.