This invention generally relates to strain sensor devices, and more specifically relates to methods for calibrating strain sensor devices incorporated into aircraft.
Structural health monitoring programs for modern aircraft typically use permanently installed structural health sensors (e.g., “strain sensors”) to compute in-flight loads on individual fleet aircraft. The load histories are used by structural fatigue life tracking methods to predict remaining structural life for individual aircraft. It has been seen that strain sensor readings for a given load can vary significantly from aircraft to aircraft due to manufacturing and installation variations. For example, strain reading errors of 10% can lead to errors of more than 50% in predicted fatigue life, thereby posing a serious obstacle to ensuring safety and minimizing maintenance costs. Hence, it is advantageous to resolve discrepancies in strain sensor readings in fatigue life tracking programs.
Known methods for calibrating aircraft strain sensors are undesirable in a number of respects. Specifically, they are typically costly, inaccurate, and provide poor repeatability, leading to substantial errors in the prediction of remaining life. One conventional approach requires that each aircraft be placed in a full-scale test rig. This approach is expensive and time-consuming for most aircraft fleets. An alternate method, using in-flight calibration, compares the strain sensor output to the loads assumed during tightly prescribed flight maneuvers. This method is considerably less accurate because maneuvers that can repeat loads on certain portions of the airframe, such as the vertical tail or canopy sill, are difficult to prescribe.
Accordingly, there is a need for improved systems and methods for calibrating structural health sensors incorporated into aircraft.
The present invention will hereinafter be described in conjunction with the appended drawings, where like designations denote like elements, and:
The following detailed description presents a number of example embodiments and is not intended to limit the invention or the application and uses of the invention. Furthermore, there is no intention to be bound by any expressed or implied theory presented in the preceding technical field, background, brief summary, or the following detailed description.
In general, the present invention relates to low-cost systems and methods for quickly calibrating a plurality of structural health sensors (or simply “sensors”) incorporated into an aircraft. As will be shown, what is provided is a flexible, automated calibration system that is portable, low-cost, can be used by personnel with relatively little training, and can be owned and operated on the fleet level.
Data acquisition subsystem 110 includes any suitable combination of hardware and software (e.g., software stored within storage 108 and executable by CPU 106) configured to acquire stress, strain, and/or force information from integrated health sensors (or simply “sensors”) 102 within aircraft 100 as well as force sensor 116 of actuation system 114. In this regard, data acquisition subsystem 110 might be coupled to sensors 102 through a pre-existing coupling or interconnect that is already provided within aircraft 100. With brief reference to
Referring again to
Calibration system 200 might include additional components which, in the interest of simplicity, have not been illustrated. For example, system 200 might include a fail-safe microcontroller or an interface to an “emergency stop” component. System 200 might also typically include a communication bus and/or a power distribution bus.
Calibration system 200 preferably includes a software environment configured to carry out the calibration methods described herein, e.g. via the use of interpreted and/or compiled software code used in connection with one or more math/science libraries, e.g., MATLAB, Numpy/SciPy, R, or the like. Thus, storage 108 may include computer-readable instructions adapted to cause CPU 106 to perform the various steps described herein.
In one embodiment, an initialization file is provided within storage 108 to define operating parameters, for example, operating parameters corresponding to each mechanical coupling point 104. Example operating parameters include actuator speed (e.g., the speed at which actuator 115 should move during testing), not-to-exceed load (e.g., the maximum force actuation system 114 should apply, as determined by sensor 116), target calibration loads (e.g., at sensor 116), and calibration load duration.
All or a portion of the components depicted in
In an alternate embodiment, actuation system 114 of
Having thus given an overview of a calibration system in accordance with various embodiments,
Next, in step 304, one or more of the plurality of mechanical coupling points 104 on aircraft 100 are selected or determined based on one or more structural characteristics of the aircraft and the location of each of the plurality of structural health sensors 102. These mechanical coupling points may be determined empirically, through computer simulation (e.g., CAD, finite element), or through any other suitable means. In general, the location of each mechanical coupling point is chosen to be relatively close to a corresponding sensor location while at the same time being amenable to mechanical coupling via a suitable adapter, as may be the case with various “hard points” on the aircraft—i.e., any location on the aircraft that is not likely to incur damage due to a localized load applied during the calibration procedure. This can be seen, conceptually, in
During the calibration process, for each of the selected mechanical coupling points, the system is configured to couple an actuation system to the mechanical coupling point via an associated one of the adapters (step 310), then apply a force to the mechanical coupling point (step 312). In one embodiment, the force is applied to the mechanical coupling points quasi-statically—that is, very slowly (e.g., about 0.1-0.5 mm/s). In a particular embodiment, the force is applied upward and substantially normal to a platform on which the aircraft rests (e.g., via a hydraulic jack subsystem or electromechanical actuator integrated into actuation system 114).
While applying a force in order to exercise the structural health sensors 102, the system acquires a force signal indicative of the force applied to the mechanical coupling point 104 and at least one structural health signal indicative of the output of one or more of the plurality of structural health sensors 102. In this way, the sensitivity of the structural health sensors to an external load may be determined. Optionally, a deflection signal indicative of the deflection experienced by the mechanical coupling point 104 may also be acquired. Further, the various signals may be filtered or otherwise conditioned as is known in the art.
Finally, calibration settings (e.g., calibration factors, sensitivity ratios, and/or offsets) are determined for the plurality of structural health sensors based on the force signals, the deflection signals, and the structural health signals associated with each of the plurality of mechanical coupling points (step 316). The sensitivity factors may be compared to analytically determined sensitivity factors (e.g., obtained from a finite-element analysis) or to experimentally determined sensitivity factors (e.g., obtained from testing in a full-scale rig). The ratio of these factors provides a set of calibration settings. The calibration settings are therefore a measure of variation with respect to the reference sensitivity of that sensor location and a given load application method. Sensitivity factors may be expressed, for example, as microstrain (strain×106) per pound of applied force (με/lbf). Calibration ratios may be then be expressed as a unitless value that scales (or “corrects”) the data determined during aircraft operation.
Depending upon the nature of the aircraft 100 under test, the number and type of mechanical adapters used during a particular testing session may vary greatly. A number of example mechanical adapters are illustrated in
In general, adapter 601 is configured to be removeably attached to the inner wing of an aircraft (i.e., at a jack point beneath the inner wing), adapter 602 is configured to be removeably attached to a horizontal stabilator (or “tail”) of an aircraft (e.g., at the spindle, a known structure), adapter 603 is configured to be removeably attached to an aircraft's outer wing tip (e.g., the underside of the wing tip), and adapter 604 is configured to be removeably attached to a vertical stabilizer.
In one embodiment, force is applied laterally and substantially parallel to the platform on which the aircraft rests by means of a self-reacting load between paired structures, (such as vertical tails), or by means of an external reaction structure corresponding to actuation system 114. In one embodiment, for example, the vertical stabilizers are calibrated by pulling them towards each other with a turnbuckle and threaded rod, wherein a load cell is installed in series with this assembly. The assembly is attached to the flap hinges after the control surfaces are removed. An electromechanical actuator may also be employed to draw the stabilizers together. Such a system is attached to the flap hinges of the control surface (without removing the control surfaces) of the stabilizer with a claw-like or other such subassembly. In general, this embodiment might employ any suitable wire, chain, rod, or the like, along with a system adapted to either pull the stabilizers together or to/from an external reaction structure (while affording a means to measure the applied force).
In summary, the systems and methods described above provide a way to determine calibration factors (e.g. sensitivities) used to account for differences in the sensor installation and aircraft build so that engineers may accurately derive in-flight loads from measured in-flight strain data (i.e., errors of less than 2% in the derived loads) for individual strain gages on fleet aircraft. These in-flight loads are used in fatigue calculations, and so the increase in accuracy significantly improves remaining life predictions, improves safety, and reduces maintenance costs and aircraft downtime.
In one embodiment, the system is provided as a portable hardware package, with integrated software, that is used to capture aircraft-to-aircraft variations in installed strain sensors, allowing accurate prediction of remaining structural life for individual fleet aircraft. The calibration factor obtained during system operation is then used by fatigue life prediction engineers in conjunction with in-flight strain readings to accurately predict remaining life and schedule maintenance for each aircraft. Through the use of an automated calibration process, the illustrated system can be used by personnel with relatively little training, easing logistical burdens for performing strain sensor calibration. Further minimizing the calibration process logistics footprint, the technology is insensitive to variations in aircraft configuration parameters such as fuel weight, tire pressure, and store configuration. The system can be customized to different aircraft or strain sensor configurations as needed. Due to portability and low cost, the technology can be owned and operated on the fleet or squadron level, providing flexibility as to when and where calibration procedures can be performed.
The embodiments and examples set forth herein were presented in order to best explain the present invention and its particular application and to thereby enable those skilled in the art to make and use the invention. However, those skilled in the art will recognize that the foregoing description and examples have been presented for the purposes of illustration and example only. The description as set forth is not intended to be exhaustive or to limit the invention to the precise form disclosed.
This application claims priority to U.S. Prov. Pat. App. No. 61/655,798, filed Jun. 5, 2012, the contents of which are hereby incorporated by reference.
This invention was made with U.S. Government support under contract number N68335-10-C-0187, awarded by the U.S. Navy. The U.S. Government has certain rights in the invention.
Number | Date | Country | |
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61655798 | Jun 2012 | US |