Methods and systems for structural health monitoring

Patent metadata

Field	Value
Publication number	US-9506836-B2
Application number	US-201213647935-A
Country	US
Kind code	B2
Filing date	Oct 9, 2012
Priority date	Oct 9, 2012
Publication date	Nov 29, 2016
Grant date	Nov 29, 2016

Abstract

Methods and apparatus for structural health monitoring are described. In one example, a method for use in designing a structural health monitoring (SHM) system for use in monitoring a host structure is described. The method includes one or more of a process for designing SHM systems for any given piece of structural hardware, a process for evaluating a given SHM system, a method to quantify the performance of a given SHM system in comparison to current inspection processes, a finite element modeling approach to determining excitation frequencies to detect damage and for selecting the best time window to use for sensed excitation signals, a Bayesian Network based data fusion technique that fuses in environmental information (load cycles induced on the structure) with a damage index (DI) to produce crack detection and estimation of crack length, and a damage location and sensor selection technique.

First claim

What is claimed is:

1. A method for measuring structural health on hot spots of an aircraft structure using a series of adjacently located actuators in the aircraft structure including at least one group, the method comprising:
generating, using the at least one group of the series of adjacently located actuators in the aircraft structure, a vibration in the aircraft structure;
collecting, by a plurality of sensors disposed in one or more testing areas of the aircraft structure, the vibration produced by the at least one group of the series of adjacently located actuators, the plurality of sensors locally disposed in one or more testing areas of the aircraft structure;
comparing, by a controller, the vibration collected by each sensor of the plurality of sensors to a reference signal for the at least one group of the series of adjacently located actuators;
generating a summation of damage indexes for the at least one group of the series of adjacently located actuators; and
using Bayesian network based data fusion of environmental information fused with the damage indexes from the at least one group of the series of adjacently located actuators to estimate a crack length in the aircraft structure based on a return signal response.

2. The method of claim 1 , further comprising using center of mass calculations for the aircraft structure to determine a location of one or more damaged areas or sections that require repair of the aircraft structure.

3. A SHM system in accordance with claim 2 , wherein said controller is further configured to output an indication that damage has occurred, the expected value of the damage characteristic, and the ninety-five percent prediction interval when the expected value of the damage characteristic exceeds the damage threshold.

4. The method of claim 1 , further comprising using Bayesian network based data fusion of load cycles induced on the aircraft structure fused with the damage indexes from the at least one group of the series of adjacently located actuators to estimate a crack length in the aircraft structure based on a return signal response.

5. The method of claim 4 , further comprising measuring, storing, and comparing the crack length periodically and recording the damage index associated with the crack length.

6. The method of claim 1 , further comprising receiving, by the controller, an indication of a value of an aging factor of the aircraft structure from at least one aging factor sensor.

7. A structural health monitoring (SHM) system for use in monitoring a host structure that may be susceptible to damage, the damage having a damage characteristic influenced, at least in part, by an aging factor, said SHM system comprising:
at least one actuator configured to couple to the host structure and propagate a vibration signal through the host structure;
at least one sensor configured to couple to the host structure, detect the vibration signal propagated through the host structure by said at least one actuator, and generate at least one signal in response to the vibration signal propagated through the host structure by said at least one actuator; and
a controller coupled to said at least one actuator and said at least one sensor, said controller configured to:
control said at least one actuator to propagate the vibration signal through the host structure;
receive the at least one vibration signal from said at least one sensor;
calculate a damage index (DI) based, at least in part, on the at least one signal;
input the calculated DI to a Bayesian network configured to output an expected value of the damage characteristic; and
determine a ninety-five percent prediction interval for the expected value of the damage characteristic when the expected value of the damage characteristic exceeds a damage threshold.

8. A SHM system in accordance with claim 7 , further comprising at least one aging factor sensor coupled to the controller and configured to provide an indication of a value of the aging factor to said controller.

9. A SHM system in accordance with claim 8 , wherein said controller is configured to input the calculated DI and the value of the aging factor to the Bayesian network configured to output the expected value of the damage characteristic.

10. A SHM system in accordance with claim 7 , wherein said controller is configured to use center of mass calculations for the aircraft structure to determine a location of one or more damaged areas or sections that require repair of the aircraft structure.

11. A SHM system in accordance with claim 7 , wherein said controller is configured to measure, store, and compare the damage characteristic periodically and recording the DI associated with the damage characteristic.

12. A SHM system in accordance with claim 7 , wherein said controller is configured to output an indication that damage has occurred, the expected value of the damage characteristic, and the ninety-five percent prediction interval when the expected value of the damage characteristic exceeds the damage threshold.

13. A method for measuring structural health on hot spots of an aircraft structure using a series of adjacently located actuators in the aircraft structure including at least one group, the method comprising:
generating, using the at least one group of the series of adjacently located actuators in the aircraft structure, a vibration in the aircraft structure;
collecting, by a plurality of sensors disposed in one or more testing areas of the aircraft structure, the vibration produced by the at least one group of the series of adjacently located actuators, the plurality of sensors locally disposed in one or more testing areas of the aircraft structure;
comparing, by a controller, the vibration collected by each sensor of the plurality of sensors to a reference signal for the at least one group of the series of adjacently located actuators;
generating a summation of damage indexes for the at least one group of the series of adjacently located actuators; and
using Bayesian network based data fusion of load cycles induced on the aircraft structure fused with the damage indexes from the at least one group of the series of adjacently located actuators to estimate a crack length in the aircraft structure based on a return signal response.

14. The method of claim 13 , further comprising measuring, storing, and comparing the crack length periodically and recording the damage index associated with the crack length.

15. The method of claim 13 , further comprising using center of mass calculations for the aircraft structure to determine a location of one or more damaged areas or sections that require repair of the aircraft structure.

16. The method of claim 13 , further comprising determining, by the controller, whether the estimated crack length exceeds a damage threshold.

17. The method of claim 16 , further comprising determining, by the controller, a ninety-five percent prediction interval for the estimate of the crack length when the estimated crack length exceeds the damage threshold.

18. The method of claim 17 , further comprising outputting, by the controller, an indication that damage has occurred, the estimated crack length, and the ninety-five percent prediction interval when the estimated crack length exceeds the damage threshold.

19. The method of claim 13 , further comprising receiving, by the controller, an indication of a value of an aging factor of the aircraft structure from at least one aging factor sensor.

20. The method of claim 19 , wher

Assignees

• Boeing Co

Inventors

• Pado Lawrence Eric
• Ihn Jeong-Beom

Classifications

• G06N7/01

  Probabilistic graphical models, e.g. probabilistic networks · CPC title

• G01M5/0033 Primary

  by determining damage, crack or wear · CPC title

• G01M5/0066 Primary

  by exciting or detecting vibration or acceleration (vibration testing of structures G01M7/00) · CPC title

• G06F30/20

  Design optimisation, verification or simulation (optimisation, verification or simulation of circuit designs G06F30/30) · CPC title

• G06F30/23 Primary

  using finite element methods [FEM] or finite difference methods [FDM] · CPC title