Patent Publication Number: US-8113866-B2

Title: Aircraft electrical connector with differential engagement and operational retention forces

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
     This application is a continuation of U.S. patent application Ser. No. 12/645,451, entitled “Aircraft Electrical Connector with Differential Engagement and Operational Retention Forces” filed on Dec. 22, 2009, which issued as U.S. Pat. No. 7,980,875 on Jul. 19, 2011, which is a continuation-in-part of U.S. patent application Ser. No. 11/681,674, entitled “Aircraft Power Connector with Differential Engagement and Operational Retention Forces” filed on Mar. 2, 2007, which issued as U.S. Pat. No. 7,871,282 on Jan. 18, 2011, which claims priority to U.S. Provisional Application No. 60/781,842, filed on Mar. 13, 2006, all of which are hereby incorporated by reference in their entirety. 
    
    
     FIELD OF THE INVENTION 
     The present invention relates generally to aircraft electrical connectors. Specifically, embodiments are disclosed wherein an aircraft power connector has differential engagement and retention forces. 
     BACKGROUND OF THE INVENTION 
     This section is intended to introduce the reader to various aspects of art that may be related to various aspects of the present system and techniques, which are described and/or claimed below. This discussion is believed to be helpful in providing the reader with background information to facilitate a better understanding of the various aspects of the present disclosure. Accordingly, it should be understood that these statements are to be read in this light, and not as admissions of prior art. 
     When an aircraft (e.g., a military aircraft or a commercial airliner) is being serviced, a stationary power system (e.g., bridge mounted power system), a fixed central power system, or a mobile ground power cart may supply electrical power necessary for basic operations while the aircraft&#39;s engines are not being used to power the aircraft. The power source may include an electrical generator (e.g., diesel or gasoline engine driven generator) or an electrical power grid. Typically, the aircraft is electrically connected to the ground power by way of an electrical connector mating. Existing ground power connectors typically include open orifices through which the connectors on the electrical aircraft are connected. The repeated connection and disconnection associated with connecting the ground power with the aircraft may wear the connectors, effectively limiting the number of connections that may be made between the aircraft and ground power. Furthermore, due to the construction of the connectors, the force needed to connect the ground power with the aircraft is often equal to the force of retention, which may create difficulties in situations where an operator may not be able to exert the requisite amount of force needed for connection and disconnection. 
     SUMMARY OF THE INVENTION 
     A system is provided for powering an aircraft while in service. The system may contain, among other features, an aircraft electrical connector containing a first electrical connector, a trigger configured to move the first electrical connector between a first position having a first retention force and a second position having a second retention force. The second retention force may be lower than the first so as to allow an operator to easily connect and disconnect the connector from the aircraft. 
     A system is provided containing an aircraft electrical connector including a first electrical connector and a biasing mechanism configured to move the first electrical connector in a first direction crosswise relative to a connection axis of the aircraft electrical connector. A trigger is coupled to the biasing mechanism. 
     A system is provided containing an aircraft electrical connector which includes, among other features, a first electrical connector configured to couple with a first mating connector, a biasing mechanism configured to move between a first position and a second position, wherein the first position has a first retention force between the first electrical connector and the first mating connector, the second position has a second retention force between the first electrical connector and the first mating connector, and the second retention force is greater than the first retention force. A trigger is coupled to the biasing mechanism. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       These and other features, aspects, and advantages of the present invention will become better understood when the following detailed description is read with reference to the accompanying drawings in which like characters represent like parts throughout the drawings, wherein: 
         FIG. 1  is a perspective view of an embodiment of an aircraft electrical connector which has been constructed in accordance with present techniques, illustrated as being disposed adjacent to a conventional onboard aircraft electrical connector; 
         FIG. 2  is a substantially top plan view of the aircraft power connector and onboard aircraft electrical connector illustrated in  FIG. 1 ; 
         FIG. 3  is a substantially side elevational view of the aircraft electrical connector and the onboard aircraft electrical connector illustrated within  FIG. 1 , wherein the two are illustrated as being electrically connected; 
         FIG. 4  is a perspective view of the aircraft electrical connector and onboard aircraft electrical connector of  FIG. 2 , illustrated in an engaged position; 
         FIG. 5  is a substantially side elevational view of the aircraft electrical connector of  FIG. 3 , illustrating one embodiment of the unique biasing mechanism in a locked position; 
         FIG. 6  is an enlarged, partial, substantially side elevational view of the aircraft electrical connector of  FIG. 5 , illustrating an embodiment of a portion of an embodiment of a biasing member in accordance with an aspect of the present techniques; 
         FIG. 7  is an enlarged, partial, substantially side elevational view of the aircraft electrical connector of  FIG. 3 , illustrating the connection of a first end portion of one of the lever arms of the biasing member of the aircraft electrical connector of  FIG. 1 , illustrated as mounted upon one end of a force-transmission cam plate member, which projects outwardly through a side wall portion of the aircraft electrical connector housing, by means of a retaining ring or snap-ring member; 
         FIG. 8  is a side elevational view of one of the substantially L-shaped lever members of one embodiment of the unique biasing mechanism of the aircraft electrical connector; 
         FIG. 9  is a top plan view of the force-transmission cam plate member of an embodiment of the unique biasing mechanism of the aircraft electrical connector; 
         FIG. 10  is an end elevational view of the force-transmission cam plate member as illustrated within  FIG. 9 ; 
         FIG. 11  is a perspective view of a retaining ring or snap-ring member used to secure together component parts of an embodiment of the unique biasing mechanism of the aircraft electrical connector; 
         FIG. 12  is a longitudinal cross-sectional view of the rotary tubular member of an embodiment of the unique biasing mechanism of the aircraft electrical connector illustrated in  FIG. 1 ; 
         FIG. 13  is a cross-sectional view of the rotary tubular member as illustrated within  FIG. 12  as taken along the lines  13 - 13  of  FIG. 12 ; 
         FIG. 14  is a longitudinal cross-sectional view of the secondary cam member of the unique biasing mechanism of the aircraft electrical connector illustrated in  FIG. 1 ; 
         FIG. 15  is a cross-sectional view of the secondary cam member as illustrated within  FIG. 14  as taken along the lines  15 - 15  of  FIG. 14 ; 
         FIG. 16  is rear perspective view of a set screw member which may be used within either one of the rotary tubular member or the secondary cam member as illustrated within  FIGS. 12 and 13 , or  FIGS. 14 and 15 , respectively; 
         FIG. 17  is a perspective view of the forward end portion of the set screw as illustrated within  FIG. 16 ; 
         FIG. 18  is a perspective view of a jam-nut member which may be utilized in conjunction with any one of the set screw members as illustrated within  FIGS. 16 and 17 ; 
         FIG. 19  is a perspective view of a plug member which may be utilized within either one of the rotary tubular member or the secondary cam member as illustrated within  FIGS. 12 and 13 , or  FIGS. 14 and 15 , respectively; 
         FIG. 20  is a perspective view of an embodiment of an aircraft electrical connector displaying certain features of the unique biasing system according to the present techniques; 
         FIG. 21  is a cross-sectional view of the aircraft electrical connector of  FIG. 20 , taken along an axial plane and displaying features consistent with the unique biasing system of the present techniques and illustrated in a disengaged position; 
         FIG. 22  is a cross-sectional view of the aircraft electrical connector of  FIG. 20 , taken along an axial plane and displaying features consistent with the unique biasing system of the present techniques and illustrated in an engaged position; 
         FIG. 23  is a cross-sectional view of the nose assembly of the aircraft electrical connector of  FIG. 21 , taken along a line  23 - 23  and illustrated in a disengaged position; 
         FIG. 24A  is a cross-sectional view of the nose assembly of the aircraft electrical connector of  FIG. 22 , taken along a line  24 - 24  and illustrated in an engaged position; 
         FIG. 24B  is an enlarged cross-sectional view of a portion of the nose assembly of the aircraft electrical connector of  FIG. 24A , illustrated in an engaged position; 
         FIG. 25  is a perspective, cross-sectional view of the nose assembly of  FIG. 24 , illustrated in an engaged position and displaying features consistent with the unique collar assembly; 
         FIG. 26  is a perspective view of the unique collar assembly, illustrated between the engaged and disengaged positions and displaying features consistent with an aspect of the present techniques; 
         FIG. 27  is a perspective view of the aircraft electrical connector and onboard aircraft electrical connector as they approach each other during operation and illustrated in a disengaged position; and 
         FIG. 28  is a perspective view of a ground support power system utilizing the unique aircraft electrical connector for powering an aircraft during servicing in accordance with an aspect of the present technique. 
     
    
    
     DETAILED DESCRIPTION OF THE ILLUSTRATED EMBODIMENTS 
     Referring now to the drawings and more particularly to  FIGS. 1-5  thereof, an embodiment of an aircraft electrical connector  10  is illustrated. The aircraft electrical connector  10 , as illustrated, contains an aircraft electrical connector housing  12 , and while the aircraft electrical connector housing  12  is illustrated as comprising a forward housing section  12 F and a rearward housing section  12 R which has an electrical cable  14  physically and electrically connected thereto, the aircraft electrical connector housing  12  may alternatively be fabricated as a one-piece construction and will effectively be treated as such for the purposes of this disclosure. More particularly, the aircraft electrical connector  10  is adapted to be physically and electrically connected to a conventional or standard onboard aircraft electrical connector  16 , which is fixedly mounted at a predetermined location upon an aircraft, so as to provide electrical power to the aircraft when the aircraft is being serviced. The onboard aircraft electrical connector  16  generally contains a mounting plate structure  18  upon which six male electrical connector pins  20  are fixedly mounted so as to project outwardly therefrom. In accordance with FCC regulations and guidelines, six male electrical connector pins  20  are arranged within two rows with each one of the two rows containing three male electrical connector pins  20 . Correspondingly, it is seen that the forward end portion of the aircraft electrical connector housing  12  is provided with six bores  22  within which six electrical connector pins, not visible in the drawings, are fixedly mounted. As with the onboard aircraft electrical connector  16 , the six bores  22  and six electrical connector pins are arranged within two rows with each one of the two rows containing three electrical connector pins. In one embodiment, the forward end portions of the six electrical connector pins that are disposed within the aircraft electrical connector housing  12  are female receptacles, and in this manner, the aircraft electrical connector  10  is able to be physically and electrically mated with the onboard aircraft electrical connector  16 . 
     It is noted that, when a conventional aircraft electrical connector is to be electrically connected to the onboard aircraft electrical connector  16 , the retention force is intentionally designed to be sufficiently large and relatively high, such as, for example, to be within the range of 80 lb±20 lb. Such a retention force may ensure that the integrity of the electrical connection will not be inadvertently adversely interrupted or otherwise compromised throughout the time when the aircraft is being serviced. This retention force is a function of, for example, the friction or interference fit defined between the external or outside diameter dimensions of the male electrical connector contact pins  20  disposed upon the onboard aircraft electrical connector  16  and the internal or inner diameter dimensions of the female receptacle portions of the electrical connector contact pins disposed within the conventional aircraft connector. 
     However, it is additionally noted that in embodiments where the retention force is sufficiently large or relatively high, the insertion force that is required to initially establish the electrical connection between the conventional aircraft electrical connector and the onboard aircraft electrical connector  16  be large or relatively high. As has been noted hereinbefore, such a relatively large or high insertion force level will sometimes present procedural problems or difficulties for operational personnel in connection with the establishment of the electrical connection between the conventional aircraft connector and the onboard aircraft electrical connector  16 . 
     In accordance with an aspect of the present techniques, the internal or inner diameter dimensions of the female receptacle portions of the electrical connector contact pins disposed within the aircraft electrical connector housing  12  are enlarged to a predetermined degree, such as, for example, one thousandth of an inch (0.001″) with respect to the external or outside diameter dimensions of the male electrical connector contact pins  20  disposed upon the onboard aircraft electrical connector  16 . In this manner, the insertion force which is required to initially mate the aircraft connector  10  with the onboard aircraft electrical connector  16 , and which is a function of, for example, the friction or interference fit, is able to be substantially reduced to a more manageable level, such as, for example, within the range of about 20 lb±5 lb, or about 15 lb±10 lb. 
     While the insertion force level characteristic of the aircraft electrical connector  10  has effectively been reduced, sufficient to assuredly retain the aircraft electrical connector  10  and the onboard aircraft electrical connector  16  physically and electrically connected to each other. Therefore, additional retention force may be provided upon the aircraft connector  10  in order to effectively raise or enhance the force level, such that subsequent to the physical and electrical connection together of the aircraft connector  10  with the onboard aircraft electrical connector  16  will assuredly be retained. 
     With reference therefore now being made to  FIGS. 1-5 , it is initially noted that the aircraft electrical connector housing  12  fabricated from a suitable rubber-type material such as, for example, neoprene rubber, polyurethane, or the like. In  FIG. 1 , a transversely or laterally extending slot  24  is formed within the forward end portion of the aircraft electrical connector housing  12  so as to extend rearwardly a predetermined distance from the front face of the aircraft electrical connector housing  12 . The slot  24  is also seen to be formed between the upper and lower rows of electrical connector bores  22  defined within the forward end portion of the aircraft connector housing  12 , and in this manner, the forward end portion of the aircraft connector housing  12  is effectively divided into upper and lower half portions. A force-transmission cam plate member  26 , as can best be seen and appreciated from  FIGS. 9 and 10 , is adapted to be inserted into the slot  24  such that the oppositely disposed end portions  28  of the force-transmission cam plate member  26  project laterally outwardly from the oppositely side wall portions of the aircraft connector housing  12 . In  FIG. 10 , it is additionally seen that the longitudinally spaced edge portions  30 ,  32  of the force-transmission cam plate member  26  have rounded or arcuate configurations so as not to abrade the rubber-type material from which the aircraft connector housing  12  is fabricated when the force-transmission cam plate member  26  is rotated. 
     In order to actuate or rotatably move the force-transmission cam plate member  26  between its first and second limit positions, a pair of lever members  34 ,  34 , each one of which has a substantially L-shaped configuration, is operatively connected to the oppositely disposed end portions  28 ,  28  of the force-transmission cam plate member  26 . More particularly, as shown in  FIG. 8 , each one of the lever members  34  has a through-slot  36  defined within a first end portion  38  thereof, while a through-bore  40  is defined within a second opposite end portion  42  of each lever member  34 . The oppositely disposed end portions  28 ,  28  of the force-transmission cam plate member  26  are adapted to be respectively inserted through the slots  36 ,  36  that are defined within the first end portions  38 ,  38  of the oppositely disposed lever members  34 ,  34  to secure the first end portions  38 ,  38  upon the oppositely disposed end portions  28 ,  28  of the force-transmission cam plate member  26 . A pair of retaining rings, snap-rings, or spring-clips  44 ,  44  (as shown in  FIGS. 7 and 11 ) are adapted to be mounted upon the oppositely disposed end portions  28 ,  28 . More particularly, as shown in  FIG. 9 , each one of the oppositely disposed end portions  28 ,  28  of the force-transmission cam plate member  26  has a pair of grooves or recesses  46 ,  48  respectively defined within the longitudinally spaced edge portions  50 ,  52  thereof. Accordingly, after the oppositely disposed end portions  28 ,  28  of the force-transmission cam plate member  26  are respectively inserted through the slots  36 ,  36  of the lever members  34 ,  34 , and when the snap-rings, retaining rings, or spring-clips  44 ,  44  are respectively snap-fitted over the oppositely disposed end portions  28 ,  28 , the retaining rings, snap-rings, or spring clips  44 ,  44  will effectively securely mount the first end portions  38 ,  38  of the lever members  34 ,  34  onto the oppositely disposed end portions  28 ,  28  of the force-transmission cam plate member  26  as shown in  FIG. 7 . 
     Continuing further, in order to actuate or rotatably move the pair of lever members  34 ,  34 , an actuating handle assembly is operatively associated with the second end portions  42 ,  42  of the lever members  34 ,  34 . More particularly, the actuating handle assembly may be a handle  54  having a substantially T-shaped configuration, a rotary member  56  rotatably mounted, around its longitudinal axis, through means of its oppositely disposed end portions being disposed within the through-bores  40 ,  40  defined within the second opposite end portions  42 ,  42  of the oppositely disposed lever members  34 ,  34 , and a secondary cam member  58  fixedly mounted upon the distal end of the handle  54 . In one embodiment, the handle  54  may contain a transversely oriented finger or hand-grasping portion  60 , and a shaft portion  62  which is adjustably mounted within the rotary member  56 . The shaft portion  62  may be fabricated, for example, from a structural member having a hexagonal cross-sectional configuration (e.g., an Allen wrench). Additionally, the upper end portion of the shaft member can be bent 90° in a first direction and then bent again, in effect back upon itself 180° in the opposite direction, so as to effectively form an integrally connected transversely oriented structural member that forms the internal cross-member of the hand-grasping portion  60 . A suitable thermoplastic material may then be molded over the upper end portion of the shaft member and the cross-member so as to form the hand-grasping portion  60 . 
     With reference being made to  FIGS. 6 ,  12 , and  13 , it is seen that the rotary member  56  may contain a hollow tubular member wherein, for example, the inner periphery thereof is internally threaded throughout its entire longitudinal or axial extent. In some embodiments, a through-bore  66  is defined within the central region of the rotary member  56  so as to permit a central portion of the shaft portion  62  of the handle  54  to pass therethrough. A pair of externally threaded set screws  68 ,  68  (illustrated in  FIGS. 16 and 17 ) are adapted to be threadedly engaged within the oppositely disposed ends of the internally threaded rotary member  56  so as to engage the shaft portion  62  of the handle  54 , and thereby fixedly secure the shaft portion  62  of the handle  54  at a particular position within the rotary member  56 . As can best be additionally seen and appreciated from  FIGS. 16 and 17 , the rear end portion of each set screw  68  has a hexagonally configured recess  70  formed therewithin so as to permit a suitable rotary driving tool, such as, for example, an Allen wrench, to be drivingly engaged with the set screw  68  in order to threadedly mount the same within one end portion of the internally threaded bore  64  of the rotary member  56 . In addition, the forward end portion of each set screw  68  is provided with a cup-shaped recess  72  such that the forwardmost point of each set screw  68  defines a linear locus having a circular or annular configuration as opposed to a solid circular surface or face. This structure enables each set screw  68  to more effectively grip one of the planar surfaces containing the hexagonally configured shaft portion  62  of the handle  54  when the set screw  68  is in fact engaged with the shaft portion  62  of the handle  54   
     Still further, in order to fixedly secure each one of the set screws  68  at its engaged position with the shaft portion  62  of the handle  54 , an externally threaded jam nut or jam set screw  74 , as illustrated within  FIG. 18 , may likewise be threadedly engaged within each one of the oppositely disposed end portions of the internally threaded bore  64  of the rotary member  56  until each one of the jam nuts or jam set screws  74 ,  74  tightly engages a respective one of the set screws  68 ,  68 . In a manner similar to that of each one of the set screws  68 , each one of the jam nuts or jam set screws  74 ,  74  has a hexagonally configured through-bore  76  defined therethrough so as to permit a suitable rotary driving tool, such as, for example, an Allen wrench, to be drivingly engaged with the jam nut or jam set screw  74  in order to respectively threadedly mount the same within one end portion of the internally threaded bore  64  of the rotary member  56 . With reference also being made to  FIGS. 1-5  and  19 , it is additionally seen that end plugs  78 ,  78 , fabricated, for example, from a suitable thermoplastic material, may be respectively inserted, in accordance with a friction or snap-fitting mode of operation, into each open end of the internally threaded bore  64  of the rotary member  56  so as to simply provide the opposite ends of the rotary member  56  with a finished appearance as well as to prevent dirt, debris, contaminants, or the like, from entering such open ends of the internally threaded bore  64 . 
     With reference being made to  FIGS. 1-6 ,  8 , and  12 , in order to respectively rotatably secure the oppositely disposed end portions of the rotary member  56  within the second end portions  42 ,  42  of the lever members  34 ,  34 , and concomitantly or conversely, in order to respectively positionally secure the second end portions  42 ,  42  of the lever members  34 ,  34  onto the oppositely disposed end portions of the rotary member  56 , it is seen, as illustrated in  FIG. 12 , that the external peripheral surface regions of each one of the oppositely disposed end portions of the rotary member  56  are provided with a pair of longitudinally or axially spaced annular recesses or grooves  80 ,  82  with a non-recessed or non-grooved region  84  defined therebetween. Accordingly, when, for example, the second end portions  42 ,  42  of the lever members  34 ,  34  are to be respectively mounted onto the end portions of the rotary member  56 , a first retaining ring, snap-ring, or spring clip  44 , is initially mounted within each one of the axially inner annular grooves or recesses  80 ,  80 . The end portions of the rotary member  56  are then respectively inserted through the through-bores  40 ,  40  such that the inner peripheral surface regions of the through-bores  40 ,  40  will respectively effectively be seated upon the external peripheral, non-recessed or non-grooved regions  84 ,  84  of the oppositely disposed end portions of the rotary member  56 . Lastly, a second retaining ring, snap-ring, or spring clip  44  is mounted within each one of the axially outer annular grooves or recesses  82 ,  82 , thereby effectively positionally trapping each one of the second end portions  42 ,  42  of the lever members  34 ,  34  upon the end portions of the rotary member  56 . These assemblies are illustrated within, for example,  FIGS. 1-4  and  6 . 
     In  FIGS. 14 and 15 , it is seen that the secondary cam member  58  is structurally similar to the rotary member  56  in that the secondary cam member  58  likewise contains a hollow tubular member wherein, for example, the inner periphery thereof is internally threaded throughout the entire longitudinal or axial extent thereof. In one particular embodiment, a blind bore  88  is formed within one centrally located side wall portion of the secondary cam member  58  so as to permit the distal end portion of the shaft portion  62  to be inserted into the blind bore  88  and effectively be seated upon the oppositely disposed internal side wall portion of the secondary cam member  58 . Subsequently, in order to fixedly secure the distal end portion of the shaft portion  62  within the secondary cam member  58 , a pair of externally threaded set screws  68 ,  68  is adapted to be threadedly engaged within the oppositely disposed ends of the internally threaded secondary cam member  58 . 
     Still further, in order to fixedly secure each one of the set screws  68  at its engaged position with the distal end portion of the shaft portion  62  of the handle  54 , an externally threaded jam nut or jam set screw  74  may be threadedly engaged within each one of the end portions of the internally threaded bore  86  of the secondary cam member  58  until each one of the jam nuts or jam set screws  74 ,  74  tightly engages a respective one of the set screws  68 ,  68 . End plugs, similar to the end plugs  78 ,  78 , as illustrated within  FIG. 19 , may be respectively inserted into each open end of the internally threaded bore  86  of the secondary cam member  58  so as to simply provide the opposite ends of the secondary cam member  58  with a finished appearance as well as to prevent dirt, debris, contaminants, or the like, from entering such open ends of the internally threaded bore  86 . 
     Having described the various structural components according to one embodiment of the aircraft electrical connector  10 , a method of operation of using the same will now be described. More particularly, when the actuating handle assembly is disposed at the position illustrated within any one of  FIGS. 1-3  whereby handle  54  has effectively been rotated in the clockwise direction, the aircraft electrical connector  10  may be disposed at its UNLOCKED position such that the secondary cam member  58  is disengaged from, or disposed out of contact with, the aircraft electrical connector housing  12 . Thus, the female receptacle portions of the electrical connector contact pins disposed within the aircraft electrical connector housing  12  may exhibit a relatively low insertion or engagement force level on the order of, for example, about 15 lb±10 lb due to the foregoing enlarged machining of the female receptacle portions of the electrical connector contact pins disposed within the aircraft electrical connector housing  12 . Accordingly, at this point in time, the aircraft electrical connector  10  can be moved by operator personnel from its disengaged position with respect to the onboard electrical connector  16 , as illustrated within  FIGS. 1 and 2 , to its position illustrated within  FIG. 3  at which the aircraft electrical connector  10  is able to be readily and easily physically mated or engaged with, and electrically connected to, the onboard aircraft electrical connector  16  in a coaxially aligned manner. 
     Subsequently, when it is desired to increase the force level defined between the aircraft electrical connector housing  12  and the onboard aircraft electrical connector  16 , the handle  54  is rotated in the counterclockwise direction around the rotary axis defined by means of the rotary member  56 , such that the secondary cam member  58  is initially moved from its disposition illustrated in  FIG. 3  to an intermediate position, as illustrated within  FIG. 4 , wherein the secondary cam member  58  is now disposed in contact with the upper surface portion of the aircraft electrical connector housing  12 . Subsequently, continued rotation of the handle  54  in the counterclockwise direction from its intermediate position, as illustrated within  FIG. 4 , to its final or LOCKED position, as illustrated within  FIG. 5 , causes the pair of lever members  34 ,  34  to undergo rotational or pivotal movement in the counterclockwise direction wherein the pair of lever members  34 ,  34  will, in turn, cause the force transmission cam plate member  26  to rotate or pivot around its longitudinal axis. 
     As mentioned, the force transmission cam plate member  26  may be disposed within the slot  24  of the aircraft electrical connector housing  12 , such that the aforenoted rotational or pivotal movement of the force transmission cam plate member  26  will effectively cause the lower half of the forward end portion of the aircraft electrical connector housing  12 , and the female receptacle portions of the electrical connector contact pins disposed within, to move downwardly a predetermined amount. This predetermined downward movement of the lower row of female receptacle portions of the electrical connector contact pins may effectively cause a predetermined amount of coaxial misalignment to be developed between the lower row of female receptacle portions of the electrical connector contact pins and the lower row of male electrical connector contact pins  20  mounted upon the onboard onboard aircraft electrical connector  16 . Accordingly, such a predetermined amount of coaxial misalignment may result in enhanced or increased surface-to-surface and frictional contact. In turn, such enhanced or increased surface-to-surface and frictional contact results in enhanced or increased retention engagement forces to be developed between the lower row of female receptacle portions of the electrical connector contact pins and the lower row of male electrical connector contact pins  20 . Accordingly, the associated disengagement resistance forces may likewise be enhanced. 
     It is to be further noted that the actuating handle assembly, containing the handle  54 , the rotary member  56 , and the secondary cam member  58 , effectively displays an over-center locking mechanism whereby when the handle  54  is rotated in the counterclockwise direction to its fully LOCKED position, as illustrated within  FIG. 5 . As such, the secondary cam member  58  will be moved slightly beyond the vertical plane within which the rotary axis, defined by means of the rotary member  56 , is located so as to effectively snap into its LOCKED position which is located at the juncture  90 . It is noted still yet further that the disposition of the handle  54  with respect to the rotary member  56  can be readily adjusted by effectively altering the particular axial location, as taken along the shaft portion  62  of the handle  54 . Altering the disposition of the handle  54  with respect to the rotary member  56  of course alters the distance or moment arm defined between the secondary cam member  58  and the rotary member  56  so as to, in turn, alter the position at which the secondary cam member  58  will in effect encounter the upper surface portion of the aircraft electrical connector housing  12 . Such an altered state or position will in turn alter the degree to which the lever members  34 ,  34 , and the attached force transmission cam plate member  26 , are rotated or pivoted before the secondary cam member  58  attains its final or LOCKED position. Accordingly, the degree to which the lower row of female receptacle portions of the electrical connector contact pins and the lower row of male electrical connector contact pins  20  are disposed in frictional contact with respect to each other can be predeterminedly adjusted. 
     It may be appreciated that when the aircraft electrical connector  10  is to be intentionally disconnected from the onboard aircraft electrical connector  16 , such as, for example, when servicing of the aircraft has been terminated, the handle  54  is rotated in the reverse, clockwise direction from its position illustrated within  FIG. 5  toward its position illustrated, for example, within  FIG. 3 . This may free or release the secondary cam member  58  from its locked position and moving the same to its released position as illustrated, for example, within  FIG. 3 . This permits the lever members  34 ,  34 , and the operatively connected force transmission cam plate member  26 , to be rotatably or pivotally moved in the clockwise direction so as to effectively relieve or reduce the force level, defined between the lower row of female receptacle portions of the electrical connector contact pins, disposed and the lower row of male electrical connector contact pins, back to its normal predetermined level of 20 lb±5 lb. The aircraft electrical connector  10  may then be easily and readily disconnected from the onboard aircraft electrical connector  16 . 
     Referring now to  FIG. 20 , one embodiment of an aircraft electrical connector  100  is illustrated depicting an implementation of a unique biasing feature. Among other features, the aircraft electrical connector  100  generally includes a nose assembly  102 , a biasing assembly  104 , and a cable assembly  106 . The cable assembly  106  may be configured to secure one or more cables  108  to the connector  100 . The cables  108  may extend through the connector  100 , such that the cable passes through the biasing assembly  104  and meets a set of large electrical connectors  110  (e.g., connector sockets) and small electrical connectors  112  (e.g., connector sockets) at an interface housing  114  contained within the biasing assembly  104 . In the depicted embodiment, the nose assembly  102  is disposed at a forward section of the connector  100  to facilitate interfacing with an aircraft, and contains the electrical connectors  110 ,  112  defined within a replaceable nose  116 . As with the previous aircraft electrical connector  12 , the electrical connectors  110 ,  112  are configured to removably interface with a mating connection on an aircraft. For example, in the depicted embodiment, the electrical connectors  110 ,  112  are female connectors configured to axially receive the male connectors  20  (e.g., pins) from an onboard aircraft electrical connector  16 . As may be appreciated, the aircraft electrical connector  100  (and thus the nose assembly  102 ) may be subjected to a number of connections and disconnections within a short period of time as a result of a large number of commercial flights (for example, at a commercial airport). Due to the repeated abutment of the forward portion of the connector  100  with the aircrafts, the replaceable nose  116  may be worn after a relatively short period of time. Thus, it may be desirable to construct the replaceable nose  112  from a robust polymeric material (e.g., impact-resistant polymeric materials) that is configured to be removably secured to the rest of the aircraft electrical connector  100  to allow an operator to replace the nose  116  as often as needed. 
     In certain embodiments, the nose assembly  102  may be disposed proximate the biasing assembly  104 , which may facilitate the biasing of the female electrical connectors  110 ,  112 . As discussed in detail below, the biasing assembly  104  may actuate crosswise movement of one or more female electrical connectors  110 ,  112  to create a lateral retention force after connection with the male connectors  20 . As illustrated, the biasing assembly  104  contains a handle  118  pivotally secured to a housing  120  by way of a pivot joint  122 . The housing  120 , in some embodiments, may be in mechanical communication with the nose assembly  102  by way of the interface housing  114 . For example, the handle  118 , when triggered, may engage a portion of the biasing assembly movably extending through the interface housing  114 . Such engagement may result in a subtle movement (e.g., crosswise) of one or more of the electrical connectors  110 ,  112 , which may either facilitate or prevent sliding of the electrical connectors  110 ,  112  over the male connectors of an aircraft, and may depend on a given implementation-specific configuration. As illustrated, the biasing assembly  104  may also contain features (e.g., electrical circuitry) configured to alert an operator as to the status of connectivity between the aircraft electrical connector  100  and the onboard aircraft electrical connector  16 . In one embodiment, the circuitry may be a simple switch configured to visually represent the current status of the connector  100 , for example, by illumination of a green or red light  124 . 
     To prevent inadvertent triggering of the biasing assembly  104  and to protect the handle  118  from accidental breakage, the aircraft electrical connector  100  may also include a handle protector  126 . The handle protector  126  may be constructed from a hard, impact-resistant polymeric material such as Kevlar®, polycarbonates, impact resistant polystyrenes, polyurethanes, and the like. Further, the handle protector  126  may have an annular region through which the cables  108  of the cable assembly  106  extend. In certain embodiments, the annular region may contain a cable seal  128  and cable seal flange  130  configured to secure and direct the cables  108  through the aircraft electrical connector  100 . It should be noted that the cable seal  128  and the cable flange  130  may have a generally annular shape, and may be adapted to receive cables  108  in specific configurations, so as to secure the cables  108  tightly to prevent inadvertent movement or disconnection. As such, the cable seal  128  and cable flange  130  may be replaceable, such that many different types of cables  108  may be used in conjunction with the aircraft electrical connector  100 . To facilitate such modularity, the handle protector  126  may be of a multiple-piece construction, such as a two-piece construction, and may be assembled by fastening two pieces of the handle protector  126  around the cable seal  128  and cable flange  130 . The two pieces that form the handle protector  126  may be secured together by any suitable securing mechanism, such as a snap-fit, interference fit, screw, or any mating connection. In the embodiment illustrated in  FIG. 20 , the two pieces are fastened together with a head cap screw inserted at a bottom receptacle  132  and a top receptacle  134  of the handle protector  126 . 
       FIG. 21  is a cross-section taken along a connection axis  136  of the aircraft electrical connector  100 , further illustrating certain features of the unique biasing assembly  104  according to one embodiment of the present technique. As depicted, in addition to the interface housing  114 , the handle  118 , the housing  120 , and the pivot joint  122 , the biasing assembly  104  also contains features configured to bias the position of the connectors  110 ,  112 . Such features may include a shaft  140 , a biasing spring  142 , a lever  144 , a shaft-lever connection  146 , and a cam shaft  148 . To protect the cables  108  which extend longitudinally (down the connection axis  136 ) through the aircraft electrical connector  100 , features which are movable may be contained within a cable protector  150 , such that the lever  144 , the shaft  140 , and other moveable parts do not abrade or come into contact with the cables  108 . 
     As depicted, the shaft  140  movably extends through a portion of the housing  120 , the interface housing  114 , and a portion of the nose assembly  102  along the connection axis  136  of the aircraft electrical connector  100 . The biasing spring  142  may be disposed circumferentially around the shaft  140  and may be constrained between one end of the interface housing  114  and a ledge region  152  of the shaft  140 , such that the shaft  140  is forwardly biased towards the nose  116 . The shaft  140  may be connected to the lever  144  at a pivot point defined by the connection  146 . In some embodiments, the connection  146  may be created between the shaft  140  and the lever  144  by a simple chain mechanism, such as a bicycle chain. The lever  144 , at one end, is connected to the handle  118  via the cam shaft  148  at the pivot point  122 . The cam shaft  148  is configured to convert the movement of the handle  118  (e.g., when the biasing assembly  104  is triggered) into a similar rotational movement of the lever  144 . In some embodiments, the cam shaft  148  may be shaped such that the handle  118 , which has an engagement area with the cam shaft  148  that is similarly shaped, may allow the direct provision of torque to the cam shaft  148  upon depression of the handle  118 , resulting in movement of the lever  144 . The movement of the lever  144  results in a concomitant rearwardly motion of the shaft  140  away from the nose  116 , resulting in the disengagement of a tapered section  154  of the shaft  140  from one or more collar protrusions  156  which abut some or all of the connectors  110 ,  112 . As such, the shaft  140  may be triggered by the motion of the handle  118 , with both the handle  118  and the shaft  140  being biased towards a resting position by the spring  142 . Accordingly, the handle  118 , the shaft  140 , and all other movable components of the biasing assembly  104  may be considered as being movable between a first and second position, the first position corresponding to depression or triggering of the handle  118  and the second position corresponding to releasing the handle  118  and opposite biasing by the spring  142 . Indeed, in some embodiments, these positions may be referred to as an open and closed position, respectively, an unlocked and locked position, respectively, or a disengaged and engaged position, respectively. Therefore, the position illustrated in  FIG. 21  may be described as a first, open, unlocked, or disengaged position. 
     Conversely,  FIG. 22  is a depiction illustrating the position of various features within the aircraft electrical connector  100  when the handle  118  has been released. That is, the biasing spring  142  is allowed to release its stored potential energy, returning the shaft  140 , the lever  144 , the shaft  148 , and the handle  118  back to their original, first, closed, locked, or engaged position. Thus, as illustrated, the shaft  140  has traveled forwardly and axially towards the nose  116 , allowing engagement of the tapered section  154  of the shaft  140  with the collar protrusions  156 , which outwardly bias the positions of some or all of the connectors  110 ,  112  due to their angle relative to the axis of the shaft  140 , as is described further below. 
     In some embodiments, the biasing spring  142  may be selected to have a specific spring constant, k, such that the force exerted by the spring (the stored potential energy of the compressed spring) is sufficient to move the various components of the biasing assembly  104  (and thus the connectors  110 ,  112 ) back to their engaged position. Such springs may be selected based on a desired retention force. For example, a spring with a higher spring constant k may create a larger retention force, as the stored potential energy of the spring  142  results in the biasing of the collar protrusions  156 . The travel of the shaft  140 , while illustrated as one embodiment displaying a particular length, may be varied as a function of a number of factors, including the number of connectors  110 ,  112  which may be engaged by the tapered section  154  of the shaft  140 , the size of the aircraft electrical connector  100 , the relative positions of the components of the biasing assembly  104 , and so forth. For example, the shaft  140  may travel only a few millimeters (e.g., between about 1 and about 40 millimeters), or may travel several inches. In other embodiments, the shaft  140  may travel between about 0.5 and about 6 inches (e.g., about 1, 1.5, 2, 3, or 4 inches). Further, the travel of the shaft  140  may be represented as a percentage traveled of the entire length of the aircraft electrical connector  100 , and may be between about 0.01 and about 10 percent of the total length of the aircraft electrical connector  100 . For example, the travel may be about 0.05, 0.1, 0.2, 0.5, 1, 1.5, 2, 3, 3.5, or 5 percent of the total length of the aircraft electrical connector. 
     Moving now to  FIG. 23 , a cross-section of the nose assembly  102  viewed down the connection axis  136  is shown, taken across a line  23 - 23  from  FIG. 21 , wherein the aircraft electrical connector is illustrated as being in the unlocked position (trigger  118  depressed). As illustrated, the cross-section of the nose assembly  102  generally includes six receptacles, which, in embodiments where the device is an aircraft electrical connector, correspond to the large electrical connectors  110  and small electrical connectors  112 . Each circular opening also includes an annular spring  170  disposed circumferentially within the connectors  110 ,  112 . The annular spring  170 , in general, is configured to maintain an electrical connection between the electrical connectors  110 ,  112  of the aircraft electrical connector  100  and the electrical connectors  20  of the aircraft. According to one aspect, the annular spring  170  is conductive and also exerts a small amount of force on the electrical connectors  20  of the aircraft to stabilize any initial engagement between the connector  100  and the aircraft. In some embodiments, such that the annular spring  170  may efficiently conduct electricity, the annular spring  170  may be a multi-lam rated at between about 10 amps and 30 amps (e.g., 20 amps). In certain of these, each annular spring  170  (multi-lam) may exert a force  172  (the total force exerted by all arrows in a single connector) on a male electrical connector  20  of the aircraft of between about 1 lbs and about 10 lbs. For example, the force exerted may be about 1, 2, 3, 4, 5, 6, 7, 8, 9, or 10 lbs of pressure. 
     According to an aspect of the present technique, the force exerted by the annular structures  170 , in total, may represent the total insertion force necessary to insert the male electrical connectors  20  into the electrical connectors  110 ,  112  of the aircraft electrical connector  100 . Thus, the total insertion force, in certain embodiments, may be between about 6 lbs and about 60 lbs. In one embodiment, the insertion force may be about 15 lbs to about 20 lbs. It is noted that, according to present embodiments, the total insertion force may be much less that what is necessary in conventional aircraft electrical connectors, which may require insertion forces up to 100 lbs. That is, the force needed for insertion may be equal to the force of retention in conventional aircraft electrical connectors, whereas the aircraft electrical connector according to the present embodiments requires a lower insertion force than what is needed or used for retention. In some embodiments, the insertion force may be less than about 20 lbs. For example, in such embodiments, the insertion force may be less than or equal to about 15, 10, 5, or 0 lbs. In other embodiments, the insertion force may be between about 10 percent and 50 percent of the retention force of a conventional aircraft connector. For example, the insertion force may be about 10, 20, 25, 30, 35, 40, 45, or 50 percent of the retention force. In another embodiment, the annular structures  170  may be eliminated. In such an embodiment, the annular structures  170  no longer exert the inward force  172  towards the center of the connectors  110 ,  112 . Of course, if the inward force  172  is eliminated, the aircraft electrical connector  100  will have a substantially zero insertion force. 
     To allow a decrease in the required insertion force, the connectors  110 ,  112  may be bored to a slightly larger diameter than what is conventionally used. Surprisingly, by slightly increasing the size of the electrical connectors  110 ,  112  (e.g., by 0.001 inches), the male connectors  20  of the aircraft may more easily slide into the six openings, avoiding scraping and loss of material, which is a common problem with conventional connectors. Of course, due to such scraping and loss of material, the number of connections that a conventional aircraft electrical connector may be able to perform may be limited to about 50 to about 200 insertions across the life a conventional aircraft electrical connector. In contrast, by enlarging the connectors  110 ,  112 , even to a small extent, the life of the aircraft electrical connector  100  may be, for example, between about 1500 and 2500 (e.g., 2000) insertions. In some embodiments, the longer lifetime of the aircraft electrical connector may be represented as a percentage relative to conventional aircraft electrical connectors. For example, the aircraft electrical connector  100  may have a lifetime, represented by the number of retained insertions, that is between about 300 percent and about fifteen hundred percent greater than that of a conventional aircraft electrical connector (e.g., about 1000 percent greater or about ten times greater). 
     Further depicted in  FIG. 23  is a collar assembly  174 . The collar assembly  174  generally includes collars  176 , which are disposed circumferentially around one or more of the connectors  110 ,  112 ; and the aforementioned collar protrusions  156 , which abut with the tapered portion  154  of the shaft  140  during biasing. In some embodiments, the collars  176  are disposed circumferentially about four of the six total electrical connectors  110 ,  112 . Such a configuration may allow biasing of the four of the six connectors  110 ,  112  from an area  178  disposed substantially centrally between the four of the six electrical connectors  110 ,  112 . The area  178  may be a circular area defined by four quarter-circle end areas  180  of the collar protrusions  156 . Generally, the area  178  is where the shaft  140  (more precisely, the shaft taper  154 ) extends forwardly and axially into the nose assembly  102 . Therefore, during operation and when the handle  118  is depressed, the shaft  140  moves rearwardly from the area  178  along the connection axis  136  of the connector  100 , causing the collar protrusions  156  to cease to be abutted by the shaft  140  as shown in  FIG. 23 . Accordingly, the collars  178 , collar protrusions  156 , end areas  180 , and the four biased connectors  110 ,  112  may move in a radially inward or crosswise direction (e.g., radially converging relationship) relative to the connection axis  136  of the connector  100  to a disengaged position as the shaft taper  154  moves out of the area  118  as shown in  FIG. 23 . In other words, the trigger  118  is depressed to cause rearward movement of the shaft  140 , and the collars  176  and the four connectors  110 ,  112  move crosswise toward one another in the radially converging relationship. For example, the body of the nose assembly  102  may provide some degree of resiliency, which biases the connectors  110 ,  112  back to a normal position when the shaft  140  is moved rearward. In this position, the connectors  110 ,  112  may be spaced similar to the male connectors  20  to enable easy insertion. 
     Referring now to  FIG. 24A , a cross-section viewed down the connection axis  136  of the nose assembly  102  is shown, taken across a line  24 - 24  from  FIG. 22 , wherein the aircraft electrical connector is illustrated as being in the locked position (trigger  118  released). As illustrated, the tapered portion  154  of the shaft  140  is in abutment with the quarter circle areas  180  of the collar protrusions  156 . The taper of the shaft  140  is configured such that the shaft  140  is thinner at the end that enters into the nose assembly  102 . During operation, as the handle  118  is released and the shaft taper  154  moves into the area  178 , and the gradual increase in diameter of the shaft  140  causes the collar protrusions  156  to move radially outward, in a crosswise direction (e.g., radially diverging relationship) relative to the connection axis  136  of the connector  100 . In such an embodiment, the biasing assembly  104  could be considered as being engaged. 
     As the biasing assembly  104  begins to be engaged, the collar protrusions  156  cause the collars  176  (and thus the positionally biased electrical connectors  110 ,  112 ) to move in a radially diverging manner, exerting a force  190  on the male electrical connectors  20  of the aircraft in a crosswise (perpendicular) relation to the longitudinal axis of the male electrical connectors  20 , which is generally parallel to the connection axis  136 . When the biasing assembly  104  is fully engaged (i.e., the shaft  140  has been fully abutted against the collar protrusions  156  and the spring  142  has been fully released), the force exerted on the male electrical connectors  20  may be between about 10 lbs and about 20 lbs per connector (e.g., about 15 lbs). In the illustrated embodiment, the biasing assembly  104  biases four of the six electrical connectors  110 ,  112 . However, in other embodiments, less or more than four connectors  110 ,  112  may be biased, as described below. In one embodiment, the sum of all forces exerted on the male electrical connectors  20  as a result of the biasing assembly  104  and the annular structures  170  (the sum force exerted on all six male electrical connectors  20 ) may be considered the overall retention force. In some embodiments, the overall retention force may be between about 60 lbs and about 100 lbs (e.g., about 80 lbs±20 lbs). 
     It should be noted that while the biasing of the connectors  110 ,  112  is performed using collars  176 , that any method of reversibly providing a force to a connector, such as connectors  110 ,  112 , and  20  in a perpendicular direction relative to a longitudinal axis (such as connection axis  136 ) of the connector to give differential retention and insertion forces is also contemplated. Such forces may include providing a lateral force (e.g., crosswise) on one or more of the male electrical connectors  20  (e.g., pins), for example forces  190 . For example, the lateral force may include squeezing, clasping, gripping, pushing, pressing, or compressing a single male electrical connector  20 , either directly or indirectly through the female connector  110 ,  112  (e.g., connector sockets). By further example, the lateral force may include squeezing, clasping, gripping, pushing, pressing, or compressing a plurality of the male electrical connectors  20 , either directly or indirectly through the female connector  110 ,  112 . As another example, the lateral force may include squeezing, clasping, gripping, pushing, pressing, or compressing at least one of the male electrical connectors  20 , either directly or indirectly through the female connector  110 ,  112 , relative to at least one or more other male connectors  20 . The lateral forces may cause movement of the male connectors  20  toward or away from one another, or the lateral forces may bias one or more male connectors  20  without causing any substantial movement of the male connectors  20 . 
     Further, if the retention force is not a result of biasing of multiple electrical connectors  110 ,  112 , then the total retention force may arise from providing a force to a single connector  20 , such that the total retention force on the single connector  20  is approximately 80 lbs±20 lbs, or may arise from providing forces to multiple connectors, such as two, three, four, five, or six connectors  20 . Nevertheless, the sum retention force, according to present embodiments, may be approximately 80 lbs±20 lbs. Likewise, if the retention force does result from connector movement, then the retention force may be provided as the biasing of two, three, four, five, or six connectors  110 ,  112  in relation to one another, with the overall retention force being approximately 80 lbs±20 lbs. In some embodiments, the provision of forces using the approaches described herein may allow a connector, such as connector  100 , to maintain a retention force of approximately 80 lbs±20 lbs after 500, 1000, 1500 or 2000 connections. However, it should be understood that various embodiments may employ different ranges of retention forces, different numbers and configurations of connectors, and so forth. 
       FIG. 24B  is an expanded view of  FIG. 24A  illustrating the directional movement of the collar protrusions  156  during engagement and disengagement of the biasing assembly  104 . In the illustrated embodiment, an outward direction  182  and an inward direction  184  are depicted, which result from abutment of the tapered section  154  against the collar protrusions of connectors  110 ,  112 . For example, when the biasing assembly  104  is engaged, the tapered section  154  abuts against collar protrusions  156 , causing lateral movement of the collar protrusions  156  (and thus, the connectors  110 ,  112 ) in the outward, radially diverging direction  182 . Conversely, when the biasing assembly is disengaged, for example when the handle  118  is depressed, the collar protrusions  156  and thus the connectors  110 ,  112  move in the converging, radially inward direction  184 . It should be noted that when the collar protrusions  156  move in the outward direction  182 , that the connectors  110 ,  112  may abut directly against the male connectors  20 , leading to a higher retention force than when the collar protrusions  156  move in the inward direction  184 , which may lead to a substantial alignment of the connectors  110 ,  112  with the male connectors  20 . 
     Moving now to  FIG. 25 , a perspective view of the cross section shown in  FIG. 24  is illustrated. As depicted, the perspective view shows the configuration of the electrical connectors  110 ,  112  which include, among other features, the inner, circumferentially-disposed annular springs  170 . The annular springs  170 , as depicted, are multi-lams displaying a striated structure which generally bow in towards the center of each connector  110 ,  112 . This bow contributes to the forces  172  which define the overall insertion force needed for the aircraft electrical connector  100 . Further, in embodiments where the annular springs  170  are coiled and protrude towards the center of each connector  110 ,  112 , the friction between the springs  170  and the male electrical connectors  20  of the aircraft may also contribute to the overall insertion force required. Indeed, the annular springs  170  may have many configurations, and any annular structure is contemplated wherein the structure is electrically conductive and exerts a force inwardly towards the center of each connector  110 ,  112  and against an inserted electrical connector. It should be noted, as well, that the annular springs  170  should display some level of wear resistance, due to the repeated movement of the connectors  110 ,  112  and their constant abutment against the male electrical connectors  20  when biasing is performed. 
     As illustrated, the collars  176  surround the biased connectors  110 ,  112  in a sleeve-like manner. Generally, the collars  176  extend from an approximately central portion of the connectors  110 ,  112  and out towards the connection end of the nose  116 , as is shown in  FIG. 26 . As depicted, a connector  110  has been removed from an annular opening  196  to further reveal features of the collar assembly  174 . The annular opening  196  may include features that allow the connectors  110 ,  112  to be removably secured to the interface housing  114  via a mating connection. For example, the connectors  110 ,  112  may be threadingly engaged with the interface housing  114  via a socket, such as a cam socket head cap screw disposed on a rear surface of the connectors  110 ,  112 . 
     Turning to the collar assembly  174 , the collar protrusions  156 , in some embodiments, may display a taper  198  (indicated as a change in thickness from one side to another) similar to that of the tapered section  154  of the shaft  140 . Accordingly, a surface  200  against which the tapered section  154  abuts may display an angle defined by the change in thickness of taper  198 . For example, the angle of the surface  200  may be substantially the same as the angle of the tapered section  154 . In one embodiment, the angle of the surface  200  may be defined as the angle of deviation from the connection axis  136  as measured at the forward section of the taper towards the nose  116 . Similarly, the angle of the tapered section  154  may be defined as the angle of deviation from the same, but in the opposite direction (towards the cable assembly  106 ). As mentioned, the tapered section  154  may be a slight taper, such that the abutment of the shaft  140  with the collar protrusions  156  may result in a gradual, radially outward motion of the collars  170 . For example, the tapered section  154  of the shaft  140  may have a taper of between about 0.5 percent and about 5 percent of the total diameter or circumference of the shaft  140 . In one particular embodiment, the taper of the tapered section  154  is about 1 percent. In another embodiment, the degree of the taper  198  may be measured by the angle of deviation form the connection axis  136 . In such an embodiment, the angle may be greater than 0 degrees and less than about 20 degrees. For example, the angle may be less than about 0.5, 1, 2, 3, 4, or 5 degrees. The taper  198  of the collar protrusions  156  may be slightly smaller than the tapered section  154  of the shaft  140 , such that instead of abutting against a collar protrusion  156  having a generally flat annular surface, the shaft  140  may abut against the tapered surface  200 . The configuration of the tapered shaft  140 , in combination with the tapered surface  200 , may allow the forces that result form abutment, such as the radially inward forces generated by the resistance to movement by the collars  176  and biased connectors  110 ,  112 , to be applied to a larger surface of the shaft  140  than would otherwise be feasible with alternative configurations. 
     It should be noted that the collar assembly  174 , being disposed towards the connecting end of the connectors  110 ,  112 , may allow the retention forces that result from biasing the position of the collars  176  (and thus the connectors  110 ,  112 ) against the male connectors  20  of the aircraft to be applied close to the attachment points where the male connectors  20  protrude away from the aircraft. For example, the approach of the aircraft electrical connector  100  to male connectors  20  of an aircraft during operation is shown in  FIG. 27 . As depicted, the aircraft has the onboard aircraft electrical connector  16  generally including an engagement area  210  defining the male connectors  20  and the surface  18  from which the male connectors  20  protrude. The onboard aircraft electrical connector  16  initially engages the nose assembly  102  by aligning the long axis of the male connectors  20  with the connection axis  136  of the aircraft electrical connector  100 . The male connectors  20  are then inserted into electrical connectors  110 ,  112  upon depression of the handle (trigger)  118  by an operator. When the handle  118  is released, the biasing assembly  104  acts upon the connectors  110 ,  112 ,  20  to generate the desired retention force. Again, the biasing assembly  104  forces crosswise or radial movement of one or more connectors  110 ,  112 , thereby creating crosswise forces between the female connectors  110 ,  112  and the male connectors  20 . The replaceable nose  116 , being constructed from a robust polymeric material, ensures that abutment of the connector  100  against the aircraft does not cause any damage to the surface  18  or the connectors  110 ,  112 . As mentioned, the placement of the collars  176  towards the connection end of the connectors  110 ,  112  allow the placement of retention forces close to the surface  18 . Such placement may allow a more secure retention than would be available using conventional aircraft electrical connectors, which typically apply their retention forces at the forward end (away from the surface  18 ). 
     Referring now to  FIG. 28 , an illustration of one embodiment of a ground support power system  220  that provides power from a ground power unit  222  to an aircraft  224  upon connection of the connectors  10 ,  100  to the onboard aircraft electrical connector  16  is depicted. The illustrated ground power unit  222  is a mobile vehicle having an onboard power supply, which provides power to the aircraft  224  through the power cable  14 ,  108  extending from the ground power unit  222  to the aircraft  224 . The power cable  14 ,  108  is releasably coupleable to the ground power unit  222  and the aircraft  224  at electrical connectors  10 ,  100  and  226 , respectively. Although the present techniques have been described with respect to an aircraft electrical connector ( 10 ,  100 ), the methods used herein may also be applicable to the connector  226 , which may incorporate unique aspects of the present technique, as described above with respect to differential retention and insertion forces. In operation, one or all of the electrical connectors  10 ,  100  and  226  may prevent inadvertent release via motion or tension in the power cable  14 ,  108 . However, excessive movement of the ground power unit  222  or the aircraft  224  or a critical event sensed in the ground power unit  222  or the aircraft  224  may cause the connectors  10 ,  100 , and/or  226  to release. 
     While only certain features of the invention have been illustrated and described herein, many modifications and changes will occur to those skilled in the art. It is, therefore, to be understood that the appended claims are intended to cover all such modifications and changes as fall within the true spirit of the invention.