DETAILED ACTION
Status of Claims
The following is a non-final, first office action in response to the communication filed 8/31/2023.
Claims 1-16 are currently pending and have been examined.
Priority
Applicant’s claim for the benefit of prior-filed application under 35 U.S.C. 119(e) or under 35 U.S.C. 120, 121, or 365(c) is acknowledged.
Information Disclosure Statement
Information Disclosure Statement received 8/31/2023 and 2/15/2024 have been reviewed and considered.
Notice of Pre-AIA or AIA Status
The present application, filed on or after March 16, 2013, is being examined under the first inventor to file provisions of the AIA .
Claim Rejections - 35 USC § 101
35 U.S.C. 101 reads as follows:
Whoever invents or discovers any new and useful process, machine, manufacture, or composition of matter, or any new and useful improvement thereof, may obtain a patent therefor, subject to the conditions and requirements of this title.
Claims 1-16 are rejected under 35 U.S.C. 101 because the claimed invention is directed to a judicial exception without significantly more.
Step 1 of the Subject Matter Eligibility Test entails considering whether the claimed subject matter falls within the four statutory categories of patentable subject matter identified by 35 U.S.C. 101: Process, machine, manufacture, or composition of matter.
Claims 1-16 are directed to a methods (process). As such, the claims are directed to statutory categories of invention.
If the claim recites a statutory category of invention, the claim requires further analysis in Step 2A. Step 2A of the Subject Matter Eligibility Test is a two-prong inquiry. In Prong One, examiners evaluate whether the claim recites a judicial exception.
Claim 1 recites abstract limitations, including those presented in bold below:
A method of designing an aircraft part comprising: simulating the aircraft part to be designed with a finite element method analysis model consisting of elements; and creating design information for the aircraft part by determining strengths of the elements by finite element method analysis targeting the finite element method analysis model, each of the strengths of the elements being within an allowable range even when any of possible loads is applied to each of the elements, wherein the possible loads are represented by combinations of three component forces by using the finite element method analysis model in which the elements are arrayed in two-dimensional directions, and wherein whether the each of the strengths of the elements is within the allowable range is confirmed only in cases where the combinations of the three component forces are critical combinations.
These limitations, as drafted, are a process that, under its broadest reasonable interpretation, cover performance of the limitations in the mind, or by a human using pen and paper, and therefore recite mental processes (e.g., simulating an aircraft part, creating design information, etc.). More specifically, as there is no recitation of a processing structure, nothing in the claim element precludes the aforementioned steps from practically being performed in the human mind, or by a human using pen and paper. The mere recitation of a generic computer does not take the claim out of the mental process grouping. Thus, the claim recites an abstract idea.
These limitations, as drafted, are also a process that, under its broadest reasonable interpretation, represent mathematical relationships (e.g. finite element method analysis model) and are therefore mathematical concepts. Thus, the claim recites an abstract idea.
If the claim recites a judicial exception in step 2A Prong One , the claim requires further analysis in step 2A Prong Two. In step 2A Prong Two, examiners evaluate whether the claim recites additional elements that integrate the exception into a practical application of that exception.
With respect to claims 1, the method steps of the invention lack any recitation of a machine, let alone a recitation which creates a substantial tie so as to impose meaningful limitations on the claims scope. Accordingly, the method steps can be performed entirely manually. Accordingly, there are no additional elements presented to integrate the abstract idea into a practical application.
Examiner notes that, to the extent that the finite element method implies the use of a computer, it would merely amount to generic computer implementation (i.e. “applying” the abstract idea).
If the additional elements do not integrate the exception into a practical application in step 2A Prong Two, then the claim is directed to the recited judicial exception, and requires further analysis under Step 2B to determine whether they provide an inventive concept (i.e., whether the additional elements amount to significantly more than the exception itself).
With respect to claims 1, the method steps of the invention lack any recitation of a machine, let alone a recitation which creates a substantial tie so as to impose meaningful limitations on the claims scope. Accordingly, the method steps can be performed entirely manually. Accordingly, nothing in the claim adds significantly more (i.e. an inventive concept) to the abstract idea.
As discussed above, to the extent that the finite element method implies the use of a computer, use of a computer would amount to instructions to apply the exception using a generic computing device/ Mere instructions to apply an exception using a generic computer component cannot provide an inventive concept. Use of a computer or other machinery in its ordinary capacity for economic or other tasks (e.g., to receive, store, or transmit data) or simply adding a general purpose computer or computer components after the fact to an abstract idea does not provide significantly more. See Affinity Labs v. DirecTV, 838 F.3d 1253, 1262, 120 USPQ2d 1201, 1207 (Fed. Cir. 2016) (cellular telephone); TLI Communications LLC v. AV Auto, LLC, 823 F.3d 607, 613, 118 USPQ2d 1744, 1748 (Fed. Cir. 2016) (computer server and telephone unit).
The various metrics/limitations of claims 2-4, and 6-9 merely narrow the previously recited abstract idea limitations (e.g., further characterize model parameters and design process(ing)), For the reasons described above with respect to claim 1, this judicial exception is not meaningfully integrated into a practical application, or significantly more than the abstract idea.
Claims 5 and 10-16, each recite a method for producing the aircraft part based on the design information created by the limitations addressed above, which when recited at such a high-level of generality, amounts no more than mere instructions to apply the exception. The recitation of claim limitations that attempt to cover any solution to an identified problem with no restriction on how the result is accomplished and no description of the mechanism for accomplishing the result, does not integrate a judicial exception into a practical application or provide significantly more because this type of recitation is equivalent to the words “apply it”. See Electric Power Group, LLC v. Alstom, S.A., 830 F.3d 1350, 1356, 119 USPQ2d 1739, 1743-44 (Fed. Cir. 2016); Intellectual Ventures I v. Symantec, 838 F.3d 1307, 1327, 120 USPQ2d 1353, 1366 (Fed. Cir. 2016); Internet Patents Corp. v. Active Network, Inc., 790 F.3d 1343, 1348, 115 USPQ2d 1414, 1417 (Fed. Cir. 2015).
Claim Rejections - 35 USC § 102
In the event the determination of the status of the application as subject to AIA 35 U.S.C. 102 and 103 (or as subject to pre-AIA 35 U.S.C. 102 and 103) is incorrect, any correction of the statutory basis (i.e., changing from AIA to pre-AIA ) for the rejection will not be considered a new ground of rejection if the prior art relied upon, and the rationale supporting the rejection, would be the same under either status.
The following is a quotation of the appropriate paragraphs of 35 U.S.C. 102 that form the basis for the rejections under this section made in this Office action:
A person shall be entitled to a patent unless –
(a)(1) the claimed invention was patented, described in a printed publication, or in public use, on sale, or otherwise available to the public before the effective filing date of the claimed invention.
(a)(2) the claimed invention was described in a patent issued under section 151, or in an application for patent published or deemed published under section 122(b), in which the patent or application, as the case may be, names another inventor and was effectively filed before the effective filing date of the claimed invention.
Claim(s) 1-2, 5 and 10 is/are rejected under 35 U.S.C. 102(a)(1) as being anticipated by Akiba (US 20160283648)
Akiba is directed to a method of making design information of an aircraft structural object.
Regarding claim 1, Akiba discloses:
1. A method of designing an aircraft part comprising:
simulating the aircraft part to be designed with a finite element method analysis model consisting of elements; and
Akiba [0023] FIG. 1 is a block diagram of a designing system of an aircraft structural object according to an implementation of the present invention.; Akiba [0044] In an optimization calculation of a weight of a structural object, analytical processing by the FEM (finite element method) is performed in order to estimate whether the structural object specified by design parameters has required strengths and flutter speed. ; Akiba [0081] Next, a method of making design information of an aircraft structural object using the designing system 1 for aircraft structural object will be explained.; Akiba [0082] FIG. 8 is a flow chart showing a processing flow for making design information of an aircraft structural object by the designing system 1 shown in FIG. 1.
creating design information for the aircraft part by determining strengths of the elements by finite element method analysis targeting the finite element method analysis model (Akiba [0045] That is, strengths and flutter speeds of a structural object specified by design parameters are estimated by the FEM analysis; Akiba [0086] In addition, at least one FEM model for FEM analyses which are performed in order to evaluate strengths and flutter speeds of the structural object is prepared.), each of the strengths of the elements being within an allowable range even when any of possible loads is applied to each of the elements (Akiba [0101] Then, in Step S6, whether strengths of the structural object are within allowable ranges respectively is determined. Specifically, it is determined whether a stress applied to each element of the FEM model obtained by the FEM analysis is within an allowable range specified by an acceptable value derived by adding a strength margin to a buckling load or the like. Conversely, it is determined whether each strength of the structural object specified by the design parameters is enough to the stress distribution obtained by the FEM analysis; Akiba [0102] In the strength evaluations of the structural object, a strength evaluation based on an allowable range or allowable ranges set by post buckling analysis of a composite material can be included. Specifically, when the strength evaluation of the spar 20 as shown in FIG. 7 is performed, a distribution of internal load Nxy of the spar web 21 corresponding to a height H and a sheet thickness t of the spar web 21 set as design parameters is calculated by the FEM structural analysis in Step S3.), wherein the possible loads are represented by combinations of three component forces by using the finite element method analysis model in which the elements are arrayed in two-dimensional directions (Akiba [0075] As shown in FIG. 7, a spar 20 which is one of structural parts made of composite materials has a platy spar web 21, and upper and lower chords 22. Therefore, the spar 20 has structure whose cross-sectional shape is an I shape. In addition, some stiffeners 23 are attached to the spar web 21 at a predetermined interval. Therefore, a shape of the spar 20 can be specified using the length L, the height H, and the sheet thickness t as parameters. Furthermore, a spatial coordinate system in which Z-axis represents the height direction of the spar 20, Y-axis represents the length direction of the spar 20, and X-axis represents the sheet thickness direction of the spar web 21 can be defined for an FEM analysis.; [0076] In FIG. 6, the horizontal axis shows internal loads Nxy (N/mm) of the spar web 21 on the XY plane while the vertical axis shows heights H (mm) of the spar web 21. Each region surrounded by curves shows the sheet thickness t.sub.min of the spar web 21 necessary for securing material strengths, such as the maximum principal strain and the maximum principal shear strain, required after buckling; Akiba [0092] In order to evaluate strengths of the structural object, it is necessary to obtain stresses of respective elements, under a distribution of air pressure, by an FEM analysis. In order to obtain stresses of the elements caused by air pressure in a high accuracy, it is desirable to consider deformation of the structural object, such as a wing, due to the static aeroelastic effect.; Akiba [0093] The static aeroelastic effect is the effect that the structural object is deformed as an elastic body by an aerodynamic force. Note that, an aerodynamic force is expressed by six component forces including a drag which is an air resistance in a direction of movement, a lift in an up-and-down direction, a lateral force by a crosswind, a pitch moment (pitching moment), a rolling moment (roll moment), and a yaw moment (yawing moment);), and wherein whether the each of the strengths of the elements is within the allowable range is confirmed only in cases where the combinations of the three component forces are critical combinations (Akiba [0065] Note that, a buckling load is a load which causes buckling. Allowable ranges for strengths of a material are determined as values obtained by giving strength margins to yield stresses or ultimate stresses of compression, tension, shear, surface pressure and the like of the material. Each of the strength margins to loads is generally given by multiplying the maximum load (limit load), which is applied at the time of operation of an aircraft, by a safety factor and a necessary coefficient. A value obtained by multiplying a limit load by the safety factor 1.5 is called the ultimate load, and used as the maximum load which guarantees not to break; Akiba [0106] Then, the strength evaluation of the structural object corresponding to the updated design parameters is performed again. Varying the design parameters includes not only varying design parameters for sizing structural members but also varying design parameters for specifying materials of structural members, design parameters expressing lamination structures of composite material and a design parameter expressing discernment information of an FEM model. Then, varying the design parameters and the calculation of stress distributions with considering deformation of the structural object are repeated until strengths of the structural object are determined to be not less than acceptable values respectively (stresses in respective elements of the FEM model are not more than acceptable values respectively).
Regarding claim 2, Akiba discloses the limitations of claim 1 and further discloses: wherein the cases where the combinations of the three component forces are the critical combinations are determined to cases where at least one of the three component forces becomes a critical value in at least one of the elements.
Akiba [0093] The static aeroelastic effect is the effect that the structural object is deformed as an elastic body by an aerodynamic force. Note that, an aerodynamic force is expressed by six component forces including a drag which is an air resistance in a direction of movement, a lift in an up-and-down direction, a lateral force by a crosswind, a pitch moment (pitching moment), a rolling moment (roll moment), and a yaw moment (yawing moment). (see also [0094]-[0100], e.g. [0095] As a practical method, deformation amounts of the structural object can be obtained as convergence values by a loop calculation which repeats calculation of the deformation amounts of the structural object according to a distribution of air pressure load, and calculation of the distribution of air pressure load according to the deformation amounts of the structural object.; [0097] Next, in Step S3, the FEM structural analysis is performed. Thereby, a stress distribution, arising inside the structural object due to a distribution of air pressure load, and deformation amounts of the structural object are calculated)
Akiba [0101]-[0104] Then, in Step S6, whether strengths of the structural object are within allowable ranges respectively is determined. Specifically, it is determined whether a stress applied to each element of the FEM model obtained by the FEM analysis is within an allowable range specified by an acceptable value derived by adding a strength margin to a buckling load or the like. Conversely, it is determined whether each strength of the structural object specified by the design parameters is enough to the stress distribution obtained by the FEM analysis…
Akiba [0105] In determination of Step S6, when it has been determined that at least one strength of the structural object is not within an allowable range, a value of a certain design parameter is varied into a next value again, in Step S1. (see also: Akiba [0064] Specifically, it is estimated whether evaluation parameters with regard to strengths of a structural object, such as a compressive buckling load, a shear buckling load, a load which causes a crippling fracture (local buckling), an Euler buckling load (column buckling load), and strengths of a material, become not less than values obtained by giving necessary strength margins to loads applied on the structural object, respectively. )
Akiba [0106] Then, the strength evaluation of the structural object corresponding to the updated design parameters is performed again. Varying the design parameters includes not only varying design parameters for sizing structural members but also varying design parameters for specifying materials of structural members, design parameters expressing lamination structures of composite material and a design parameter expressing discernment information of an FEM model. Then, varying the design parameters and the calculation of stress distributions with considering deformation of the structural object are repeated until strengths of the structural object are determined to be not less than acceptable values respectively (stresses in respective elements of the FEM model are not more than acceptable values respectively).
Regarding claim 5, Akiba (as shown above) discloses the limitations of claim 1 and further discloses: a method of producing the aircraft part, comprising: producing the aircraft part based on the design information created by the method of designing the aircraft part according to claim 1. (Akiba [0115] Furthermore, a combination of design parameters corresponding to the optimal value of weight of the structural object can be obtained as optimal design information of the aircraft structural object. Then, an aircraft structural object can be manufactured actually in accordance with the design information of the aircraft structural object created by the above mentioned method. An aircraft structural object manufactured by such a production method becomes a structural object whose weight has been optimized. Therefore, further weight saving of the whole aircraft as a finished product can be attained.)
Regarding claim 10, Akiba (as shown above) discloses the limitations of claim 2 and further discloses: A method of producing the aircraft part, comprising: producing the aircraft part based on the design information created by the method of designing the aircraft part according to claim 2. (Akiba [0115] Furthermore, a combination of design parameters corresponding to the optimal value of weight of the structural object can be obtained as optimal design information of the aircraft structural object. Then, an aircraft structural object can be manufactured actually in accordance with the design information of the aircraft structural object created by the above mentioned method. An aircraft structural object manufactured by such a production method becomes a structural object whose weight has been optimized. Therefore, further weight saving of the whole aircraft as a finished product can be attained).
Claim Rejections - 35 USC § 103
In the event the determination of the status of the application as subject to AIA 35 U.S.C. 102 and 103 (or as subject to pre-AIA 35 U.S.C. 102 and 103) is incorrect, any correction of the statutory basis (i.e., changing from AIA to pre-AIA ) for the rejection will not be considered a new ground of rejection if the prior art relied upon, and the rationale supporting the rejection, would be the same under either status.
The following is a quotation of 35 U.S.C. 103 which forms the basis for all obviousness rejections set forth in this Office action:
A patent for a claimed invention may not be obtained, notwithstanding that the claimed invention is not identically disclosed as set forth in section 102, if the differences between the claimed invention and the prior art are such that the claimed invention as a whole would have been obvious before the effective filing date of the claimed invention to a person having ordinary skill in the art to which the claimed invention pertains. Patentability shall not be negated by the manner in which the invention was made.
The factual inquiries for establishing a background for determining obviousness under 35 U.S.C. 103 are summarized as follows:
1. Determining the scope and contents of the prior art.
2. Ascertaining the differences between the prior art and the claims at issue.
3. Resolving the level of ordinary skill in the pertinent art.
4. Considering objective evidence present in the application indicating obviousness or nonobviousness.
Claim(s) 3, 6, 11 and 13 is/are rejected under 35 U.S.C. 103 as being unpatentable over Akiba (US 20160283648) in view of Wloch (US 20210216688).
Regarding claim 3, Akiba discloses the limitations of claim 1 and further discloses: wherein the critical combinations are determined to combinations of three component force values corresponding to points plotted … in a three-dimensional coordinate space ….
Akiba [0075] As shown in FIG. 7, a spar 20 which is one of structural parts made of composite materials has a platy spar web 21, and upper and lower chords 22. Therefore, the spar 20 has structure whose cross-sectional shape is an I shape. In addition, some stiffeners 23 are attached to the spar web 21 at a predetermined interval. Therefore, a shape of the spar 20 can be specified using the length L, the height H, and the sheet thickness t as parameters. Furthermore, a spatial coordinate system in which Z-axis represents the height direction of the spar 20, Y-axis represents the length direction of the spar 20, and X-axis represents the sheet thickness direction of the spar web 21 can be defined for an FEM analysis.; [0076] In FIG. 6, the horizontal axis shows internal loads Nxy (N/mm) of the spar web 21 on the XY plane while the vertical axis shows heights H (mm) of the spar web 21. Each region surrounded by curves shows the sheet thickness t.sub.min of the spar web 21 necessary for securing material strengths, such as the maximum principal strain and the maximum principal shear strain, required after buckling.;
Akiba [0092] In order to evaluate strengths of the structural object, it is necessary to obtain stresses of respective elements, under a distribution of air pressure, by an FEM analysis. In order to obtain stresses of the elements caused by air pressure in a high accuracy, it is desirable to consider deformation of the structural object, such as a wing, due to the static aeroelastic effect.; Akiba [0093] The static aeroelastic effect is the effect that the structural object is deformed as an elastic body by an aerodynamic force. Note that, an aerodynamic force is expressed by six component forces including a drag which is an air resistance in a direction of movement, a lift in an up-and-down direction, a lateral force by a crosswind, a pitch moment (pitching moment), a rolling moment (roll moment), and a yaw moment (yawing moment);
Akiba, as shown above, discloses the characterization of a 3-dimensional coordinate space in which at least three component force are applied. Wloch is directed to an aerodynamic simulation. Wloch more explicitly discloses that force values corresponding to points plotted on an outermost shell of a point cloud formed in a three-dimensional coordinate space of which three axes represent values of the three component forces when the combinations of the three component forces representing the possible loads are plotted in the three-dimensional coordinate space.
Wloch [0020] Referring now to FIG. 2, aspects of the present innovation will be described in which a three-dimensional computer model is segmented into a number of different portions, referred to herein as “model surface portions.” Each comprises a surface that is exposed to virtual wind, air pressure, etc. … For example, a wing of an aircraft may be broken up into perhaps 100 surface model portions, none of which would be recognizable as a wing. That said, solutions may be employed in which a single surface model portion is used for an entire wing, the fuselage, etc. Triangular and/or quadrilateral shapes will be desirable in many examples.
Wloch [0024] At the conclusion of the above workflow, simulated aerodynamic forces are available for each of the model surface portions of the three-dimensional model. Example forces on model surface portion 210c are shown in FIG. 3, indicating translational forces in X, Y and Z directions, and rotational forces about X, Y, and Z axes (pitch, yaw, roll). The existence and magnitude of these forces are based on the parameters/coefficients associated with the model surface portion. Based on a combination of simulated aerodynamic forces (i.e., a combination of the forces for each decomposed model surface portion), simulated aerodynamic performance of the three-dimensional model may be calculated. In other words, performance for the model as a whole is derived based on the forces at each surface portion.
Wloch [0028] At 402, the method entails decomposing a three-dimensional model, such as model 100, into model surface portions (e.g., portions 210), each of which has surface portion parameters (e.g., parameters 220 in FIG. 2) that influence simulated aerodynamic forces on the three-dimensional model… For example, the wing of a large aircraft may be decomposed into 20, 50 or more model surface portions. (Examiner note: each portion is associated with a 3D position, and orientation)
Wloch [0029] wide variety of parameters may be employed. Examples for parameters 220 include shape and size-related parameters, such as (1) 3D position of the surface portion; (2) 3D orientation of the surface portion; (3) front-facing and/or back facing area; and (4) local surface damage ratio. Other parameters may be directed to aerodynamic properties of the surface, such as (1) normal and tangential drag coefficients for front-facing and rear-facing aspects of the surface portion; (2) attached airflow coefficient; (3) detached airflow coefficient; (4) maximum angle of attack (stall alpha); and (5) local resistance to rotating air Cfmx, Cfmy, Cfmz. Other parameters may relate to airflow state, such as (1) current airflow attach ratio; (2) variation of current airflow attach ratio; (3) current air turbulence ratio; (4) local air speed in X, Y and Z directions; (5) local air rotation speed about X, Y and Z axes; (6) local air mach coefficient; and (7) local air density. Regarding the above factors, they may be implemented as coefficients or any other suitable way to as to influence the X/Y/Z pitch/yaw/roll force effects on surface portion 210c shown in FIG. 3. Further, these factors should be understood as non-limiting; a variety of other factors may be employed.
Wloch [0033] Continuing with 406 of method 400, the workflow includes, for each of the model surface portions, calculating a simulated aerodynamic force (or forces) based on the surface portion parameters of the model surface portion (see forces of FIG. 3). As discussed above, the surface portion parameters may include parameters relating to the shape/orientation of the model surface portion, aerodynamic properties of the model surface portion (drag coefficients, resistance to rotating air), and airflow state (air translation/rotation speed, air density); see also [0031]
One of ordinary skill in the art at the time of filing would have recognized that applying the known technique of Wloch to Akiba would have yielded predictable results and resulted in an improved system for simulating aerodynamic performance in a 3D model.
Regarding claim 6, Akiba discloses the limitations of claim 2 and further discloses: wherein the critical combinations are determined to combinations of three component force values corresponding to points plotted … in a three-dimensional coordinate space ….
Akiba [0075] As shown in FIG. 7, a spar 20 which is one of structural parts made of composite materials has a platy spar web 21, and upper and lower chords 22. Therefore, the spar 20 has structure whose cross-sectional shape is an I shape. In addition, some stiffeners 23 are attached to the spar web 21 at a predetermined interval. Therefore, a shape of the spar 20 can be specified using the length L, the height H, and the sheet thickness t as parameters. Furthermore, a spatial coordinate system in which Z-axis represents the height direction of the spar 20, Y-axis represents the length direction of the spar 20, and X-axis represents the sheet thickness direction of the spar web 21 can be defined for an FEM analysis.; [0076] In FIG. 6, the horizontal axis shows internal loads Nxy (N/mm) of the spar web 21 on the XY plane while the vertical axis shows heights H (mm) of the spar web 21. Each region surrounded by curves shows the sheet thickness t.sub.min of the spar web 21 necessary for securing material strengths, such as the maximum principal strain and the maximum principal shear strain, required after buckling.;
Akiba [0092] In order to evaluate strengths of the structural object, it is necessary to obtain stresses of respective elements, under a distribution of air pressure, by an FEM analysis. In order to obtain stresses of the elements caused by air pressure in a high accuracy, it is desirable to consider deformation of the structural object, such as a wing, due to the static aeroelastic effect.; Akiba [0093] The static aeroelastic effect is the effect that the structural object is deformed as an elastic body by an aerodynamic force. Note that, an aerodynamic force is expressed by six component forces including a drag which is an air resistance in a direction of movement, a lift in an up-and-down direction, a lateral force by a crosswind, a pitch moment (pitching moment), a rolling moment (roll moment), and a yaw moment (yawing moment); (see also [0093]-[0100])
Akiba, as shown above, discloses the characterization of a 3-dimensional coordinate space in which at least three component force are applied. Wloch is directed to an aerodynamic simulation. Wloch more explicitly discloses that force values corresponding to points plotted on an outermost shell of a point cloud formed in a three-dimensional coordinate space of which three axes represent values of the three component forces when the combinations of the three component forces representing the possible loads are plotted in the three-dimensional coordinate space.
Wloch [0020] Referring now to FIG. 2, aspects of the present innovation will be described in which a three-dimensional computer model is segmented into a number of different portions, referred to herein as “model surface portions.” Each comprises a surface that is exposed to virtual wind, air pressure, etc. … For example, a wing of an aircraft may be broken up into perhaps 100 surface model portions, none of which would be recognizable as a wing. That said, solutions may be employed in which a single surface model portion is used for an entire wing, the fuselage, etc. Triangular and/or quadrilateral shapes will be desirable in many examples.
Wloch [0024] At the conclusion of the above workflow, simulated aerodynamic forces are available for each of the model surface portions of the three-dimensional model. Example forces on model surface portion 210c are shown in FIG. 3, indicating translational forces in X, Y and Z directions, and rotational forces about X, Y, and Z axes (pitch, yaw, roll). The existence and magnitude of these forces are based on the parameters/coefficients associated with the model surface portion. Based on a combination of simulated aerodynamic forces (i.e., a combination of the forces for each decomposed model surface portion), simulated aerodynamic performance of the three-dimensional model may be calculated. In other words, performance for the model as a whole is derived based on the forces at each surface portion.
Wloch [0028] At 402, the method entails decomposing a three-dimensional model, such as model 100, into model surface portions (e.g., portions 210), each of which has surface portion parameters (e.g., parameters 220 in FIG. 2) that influence simulated aerodynamic forces on the three-dimensional model… For example, the wing of a large aircraft may be decomposed into 20, 50 or more model surface portions. (Examiner note: each portion is associated with a 3D position, and orientation)
Wloch [0029] wide variety of parameters may be employed. Examples for parameters 220 include shape and size-related parameters, such as (1) 3D position of the surface portion; (2) 3D orientation of the surface portion; (3) front-facing and/or back facing area; and (4) local surface damage ratio. Other parameters may be directed to aerodynamic properties of the surface, such as (1) normal and tangential drag coefficients for front-facing and rear-facing aspects of the surface portion; (2) attached airflow coefficient; (3) detached airflow coefficient; (4) maximum angle of attack (stall alpha); and (5) local resistance to rotating air Cfmx, Cfmy, Cfmz. Other parameters may relate to airflow state, such as (1) current airflow attach ratio; (2) variation of current airflow attach ratio; (3) current air turbulence ratio; (4) local air speed in X, Y and Z directions; (5) local air rotation speed about X, Y and Z axes; (6) local air mach coefficient; and (7) local air density. Regarding the above factors, they may be implemented as coefficients or any other suitable way to as to influence the X/Y/Z pitch/yaw/roll force effects on surface portion 210c shown in FIG. 3. Further, these factors should be understood as non-limiting; a variety of other factors may be employed.
Wloch [0033] Continuing with 406 of method 400, the workflow includes, for each of the model surface portions, calculating a simulated aerodynamic force (or forces) based on the surface portion parameters of the model surface portion (see forces of FIG. 3). As discussed above, the surface portion parameters may include parameters relating to the shape/orientation of the model surface portion, aerodynamic properties of the model surface portion (drag coefficients, resistance to rotating air), and airflow state (air translation/rotation speed, air density). See also [0031]
One of ordinary skill in the art at the time of filing would have recognized that applying the known modeling technique of Wloch to Akiba would have yielded predictable results and resulted in an improved system for simulating aerodynamic performance in a 3D model for the purpose of optimization of the surface model portion.
Regarding claim 11, the combination of Akiba and Wloch (as shown above) discloses the limitations of claim 3 and further discloses: A method of producing the aircraft part, comprising: producing the aircraft part based on the design information created by the method of designing the aircraft part according to claim 3. (Akiba [0115] Furthermore, a combination of design parameters corresponding to the optimal value of weight of the structural object can be obtained as optimal design information of the aircraft structural object. Then, an aircraft structural object can be manufactured actually in accordance with the design information of the aircraft structural object created by the above mentioned method. An aircraft structural object manufactured by such a production method becomes a structural object whose weight has been optimized. Therefore, further weight saving of the whole aircraft as a finished product can be attained)
Regarding claim 13, the combination of Akiba and Wloch (as shown above) discloses the limitations of claim 6 and further discloses: A method of producing the aircraft part, comprising: producing the aircraft part based on the design information created by the method of designing the aircraft part according to claim 6. (Akiba [0115] Furthermore, a combination of design parameters corresponding to the optimal value of weight of the structural object can be obtained as optimal design information of the aircraft structural object. Then, an aircraft structural object can be manufactured actually in accordance with the design information of the aircraft structural object created by the above mentioned method. An aircraft structural object manufactured by such a production method becomes a structural object whose weight has been optimized. Therefore, further weight saving of the whole aircraft as a finished product can be attained)
Claim(s) 4, 7, 12 and 14 is/are rejected under 35 U.S.C. 103 as being unpatentable over Akiba (US 20160283648) in view of Balabanov (US 20160193806).
Regarding claim 4, Akiba discloses the limitations of claim 1 and further discloses: wherein the aircraft part at least partially made of a fiber reinforced plastic having an anisotropic strength distribution is simulated by the finite element method analysis model, and an anisotropic strength is determined for each element simulating a portion made of the fiber reinforced plastic.
Akiba [0032] Meanwhile, when a material of a structural member is a composite material, such as GFRP (glass fiber reinforced plastics) or CFRP (carbon fiber reinforced plastics), it is preferable to set parameters for specifying a lamination structure of the composite material as design parameters, in addition to parameters for determining a size of the structural member made of the composite material, from a viewpoint of further minimizing weight of a structural object.
Akiba [0039]-[0042] generally disclosing patters of lamination structure (e.g., including setting a rate of the fiber reinforced layers 11 whose orientation angle is ±45° to be 40%, 50%, 60%, and 70%,, information for identifying a pattern of lamination structure, an inclined angle of the composite material 10 itself, etc., which can be set of design parameters
Akiba [0059] As described above, when plural FEM models are set as values of a design parameter, sizes and materials of structural members, such as a panel, spars, ribs, stringers, and a honeycomb structure, which compose one selected FEM model also become values of design parameters. Furthermore, when a material of a certain structural member is a composite material, a lamination structure of the composite material can also be expressed by a design parameter or design parameters as described above. As a matter of course, when a structural pattern has been fixed, only sizes, materials and the like of structural members may also be set as design parameters using a single FEM model.
Akiba [0067] Design, which prevents buckling from occurring at not more than the ultimate load, by evaluating strengths of a structural object is called non-buckling design. Nevertheless, when a material of a structural object is a composite material, the non-buckling design cannot fully take advantages of fiber reinforcing of the composite material. Thus, when a structural object is a composite material, it is desirable to perform post-buckling design which accepts buckling at not more than the ultimate load, from a viewpoint of attaining a further weight reduction of a structural object.
Akiba [0068] In the case of performing the post-buckling design of a composite material, it is necessary to estimate strengths of the composite material after buckling by performing a post-buckling analysis. Specifically, it is necessary to estimate whether strengths of a composite material after buckling have sufficient strength margins to loads applied on a structural object.
Akiba [0069] The post-buckling analysis is a nonlinear FEM analysis, of which input information is a stress distribution inside a structural object made of a composite material, for obtaining sheet thicknesses and sizes of the structural object, necessary for securing strengths required after buckling. The stress distribution inside the structural object made of a composite material, which is an input of the post-buckling analysis, can also be obtained as a result of an FEM analysis as described above.
Akiba [0071] Thus, allowable ranges of sheet thicknesses corresponding to necessary strengths after buckling, obtained as a result of a post-buckling analysis of a composite material, can be set as constraint conditions of the optimization calculation. That is, constraint conditions for securing the strengths of the composite material after the buckling can be set as the constraint conditions of the optimization calculation.
Akiba discloses strength distribution of fiber reinforced plastics, which very strongly suggests, but does not explicitly state a consideration for anisotropic strength distribution. Balabanov more explicitly discloses consideration of anisotropic strength distribution when modeling composite materials (Balabanov [0041] Most composite laminates are highly anisotropic. Anisotropy can be used to control dynamic mechanical behavior in a continuum. In practice, composite laminates consist of dozens to hundreds of stacked layers or plies. It is well known that mechanical behavior of individual anisotropic layers in a composite laminate can be used to model the mechanical response of the laminate. This allows designers to tailor the elastic properties and orientation of each layer (i.e., ply) so that the mechanical response of the composite laminate will be optimized; [0081] FIG. 14 shows steps of a process for designing composite skin-stringer structures that employs optimization and finite element analysis; [0086] Referring again to FIG. 14, if the finite element analysis verifies that skin-stringer delamination will be suppressed by the candidate stringer layup design, then further analyses are performed, such as notched strength, sublaminate stability, thermal residual stresses, and interpenetration. Based on the results of these analyses, a determination is made whether the candidate stringer layup design is acceptable (step 88)..)
One of ordinary skill in the art at the time of filing would have recognized that applying the known technique of Balabanov to Akiba would have yielded predictable results and resulted in an improved simulation of strength distribution based on material composition.
Regarding claim 7, Akiba discloses the limitations of claim 2 and further discloses: wherein the aircraft part at least partially made of a fiber reinforced plastic having an anisotropic strength distribution is simulated by the finite element method analysis model, and an anisotropic strength is determined for each element simulating a portion made of the fiber reinforced plastic.
Akiba [0032] Meanwhile, when a material of a structural member is a composite material, such as GFRP (glass fiber reinforced plastics) or CFRP (carbon fiber reinforced plastics), it is preferable to set parameters for specifying a lamination structure of the composite material as design parameters, in addition to parameters for determining a size of the structural member made of the composite material, from a viewpoint of further minimizing weight of a structural object.
Akiba [0039]-[0042] generally disclosing patters of lamination structure (e.g., including setting a rate of the fiber reinforced layers 11 whose orientation angle is ±45° to be 40%, 50%, 60%, and 70%,, information for identifying a pattern of lamination structure, an inclined angle of the composite material 10 itself, etc., which can be set of design parameters
Akiba [0059] As described above, when plural FEM models are set as values of a design parameter, sizes and materials of structural members, such as a panel, spars, ribs, stringers, and a honeycomb structure, which compose one selected FEM model also become values of design parameters. Furthermore, when a material of a certain structural member is a composite material, a lamination structure of the composite material can also be expressed by a design parameter or design parameters as described above. As a matter of course, when a structural pattern has been fixed, only sizes, materials and the like of structural members may also be set as design parameters using a single FEM model.
Akiba [0067] Design, which prevents buckling from occurring at not more than the ultimate load, by evaluating strengths of a structural object is called non-buckling design. Nevertheless, when a material of a structural object is a composite material, the non-buckling design cannot fully take advantages of fiber reinforcing of the composite material. Thus, when a structural object is a composite material, it is desirable to perform post-buckling design which accepts buckling at not more than the ultimate load, from a viewpoint of attaining a further weight reduction of a structural object.
Akiba [0068] In the case of performing the post-buckling design of a composite material, it is necessary to estimate strengths of the composite material after buckling by performing a post-buckling analysis. Specifically, it is necessary to estimate whether strengths of a composite material after buckling have sufficient strength margins to loads applied on a structural object.
Akiba [0069] The post-buckling analysis is a nonlinear FEM analysis, of which input information is a stress distribution inside a structural object made of a composite material, for obtaining sheet thicknesses and sizes of the structural object, necessary for securing strengths required after buckling. The stress distribution inside the structural object made of a composite material, which is an input of the post-buckling analysis, can also be obtained as a result of an FEM analysis as described above.
Akiba [0071] Thus, allowable ranges of sheet thicknesses corresponding to necessary strengths after buckling, obtained as a result of a post-buckling analysis of a composite material, can be set as constraint conditions of the optimization calculation. That is, constraint conditions for securing the strengths of the composite material after the buckling can be set as the constraint conditions of the optimization calculation.
Akiba discloses strength distribution of fiber reinforced plastics, which strongly suggests, but does not explicitly state a consideration for anisotropic strength distribution. Balabanov more explicitly discloses consideration of anisotropic strength distribution when modeling composite materials (Balabanov [0041] Most composite laminates are highly anisotropic. Anisotropy can be used to control dynamic mechanical behavior in a continuum. In practice, composite laminates consist of dozens to hundreds of stacked layers or plies. It is well known that mechanical behavior of individual anisotropic layers in a composite laminate can be used to model the mechanical response of the laminate. This allows designers to tailor the elastic properties and orientation of each layer (i.e., ply) so that the mechanical response of the composite laminate will be optimized; [0081] FIG. 14 shows steps of a process for designing composite skin-stringer structures that employs optimization and finite element analysis; [0086] Referring again to FIG. 14, if the finite element analysis verifies that skin-stringer delamination will be suppressed by the candidate stringer layup design, then further analyses are performed, such as notched strength, sublaminate stability, thermal residual stresses, and interpenetration. Based on the results of these analyses, a determination is made whether the candidate stringer layup design is acceptable (step 88)..)
One of ordinary skill in the art at the time of filing would have recognized that applying the known technique of Balabanov to Akiba would have yielded predictable results and resulted in an improved simulation of strength distribution based on material composition.
Regarding claim 12, the combination of Akiba and Balabanov (as shown above) discloses the limitations of claim 4 and further discloses: A method of producing the aircraft part, comprising: producing the aircraft part based on the design information created by the method of designing the aircraft part according to claim 4. (Akiba [0115] Furthermore, a combination of design parameters corresponding to the optimal value of weight of the structural object can be obtained as optimal design information of the aircraft structural object. Then, an aircraft structural object can be manufactured actually in accordance with the design information of the aircraft structural object created by the above mentioned method. An aircraft structural object manufactured by such a production method becomes a structural object whose weight has been optimized. Therefore, further weight saving of the whole aircraft as a finished product can be attained).
Regarding claim 14, the combination of Akiba and Balabanov (as shown above) discloses the limitations of claim 7 and further discloses: A method of producing the aircraft part, comprising: producing the aircraft part based on the design information created by the method of designing the aircraft part according to claim 7. (Akiba [0115] Furthermore, a combination of design parameters corresponding to the optimal value of weight of the structural object can be obtained as optimal design information of the aircraft structural object. Then, an aircraft structural object can be manufactured actually in accordance with the design information of the aircraft structural object created by the above mentioned method. An aircraft structural object manufactured by such a production method becomes a structural object whose weight has been optimized. Therefore, further weight saving of the whole aircraft as a finished product can be attained)
Claim(s) 8-9, and 15-16 is/are rejected under 35 U.S.C. 103 as being unpatentable over Akiba (US 20160283648) in view of Wloch (US 20210216688) and further in view of Balabanov (US 20160193806).
Regarding claim 8, the combination of Akiba and Wloch discloses the limitations of claim 3 and further discloses: wherein the aircraft part at least partially made of a fiber reinforced plastic having an anisotropic strength distribution is simulated by the finite element method analysis model, and an anisotropic strength is determined for each element simulating a portion made of the fiber reinforced plastic.
Akiba [0032] Meanwhile, when a material of a structural member is a composite material, such as GFRP (glass fiber reinforced plastics) or CFRP (carbon fiber reinforced plastics), it is preferable to set parameters for specifying a lamination structure of the composite material as design parameters, in addition to parameters for determining a size of the structural member made of the composite material, from a viewpoint of further minimizing weight of a structural object.
Akiba [0039]-[0042] generally disclosing patters of lamination structure (e.g., including setting a rate of the fiber reinforced layers 11 whose orientation angle is ±45° to be 40%, 50%, 60%, and 70%,, information for identifying a pattern of lamination structure, an inclined angle of the composite material 10 itself, etc., which can be set of design parameters
Akiba [0059] As described above, when plural FEM models are set as values of a design parameter, sizes and materials of structural members, such as a panel, spars, ribs, stringers, and a honeycomb structure, which compose one selected FEM model also become values of design parameters. Furthermore, when a material of a certain structural member is a composite material, a lamination structure of the composite material can also be expressed by a design parameter or design parameters as described above. As a matter of course, when a structural pattern has been fixed, only sizes, materials and the like of structural members may also be set as design parameters using a single FEM model.
Akiba [0067] Design, which prevents buckling from occurring at not more than the ultimate load, by evaluating strengths of a structural object is called non-buckling design. Nevertheless, when a material of a structural object is a composite material, the non-buckling design cannot fully take advantages of fiber reinforcing of the composite material. Thus, when a structural object is a composite material, it is desirable to perform post-buckling design which accepts buckling at not more than the ultimate load, from a viewpoint of attaining a further weight reduction of a structural object.
Akiba [0068] In the case of performing the post-buckling design of a composite material, it is necessary to estimate strengths of the composite material after buckling by performing a post-buckling analysis. Specifically, it is necessary to estimate whether strengths of a composite material after buckling have sufficient strength margins to loads applied on a structural object.
Akiba [0069] The post-buckling analysis is a nonlinear FEM analysis, of which input information is a stress distribution inside a structural object made of a composite material, for obtaining sheet thicknesses and sizes of the structural object, necessary for securing strengths required after buckling. The stress distribution inside the structural object made of a composite material, which is an input of the post-buckling analysis, can also be obtained as a result of an FEM analysis as described above.
Akiba [0071] Thus, allowable ranges of sheet thicknesses corresponding to necessary strengths after buckling, obtained as a result of a post-buckling analysis of a composite material, can be set as constraint conditions of the optimization calculation. That is, constraint conditions for securing the strengths of the composite material after the buckling can be set as the constraint conditions of the optimization calculation.
Akiba discloses strength distribution of fiber reinforced plastics, which strongly suggests, but does not explicitly state a consideration for anisotropic strength distribution. Balabanov more explicitly discloses consideration of anisotropic strength distribution when modeling composite materials (Balabanov [0041] Most composite laminates are highly anisotropic. Anisotropy can be used to control dynamic mechanical behavior in a continuum. In practice, composite laminates consist of dozens to hundreds of stacked layers or plies. It is well known that mechanical behavior of individual anisotropic layers in a composite laminate can be used to model the mechanical response of the laminate. This allows designers to tailor the elastic properties and orientation of each layer (i.e., ply) so that the mechanical response of the composite laminate will be optimized; [0081] FIG. 14 shows steps of a process for designing composite skin-stringer structures that employs optimization and finite element analysis; [0086] Referring again to FIG. 14, if the finite element analysis verifies that skin-stringer delamination will be suppressed by the candidate stringer layup design, then further analyses are performed, such as notched strength, sublaminate stability, thermal residual stresses, and interpenetration. Based on the results of these analyses, a determination is made whether the candidate stringer layup design is acceptable (step 88)..)
One of ordinary skill in the art at the time of filing would have recognized that applying the known technique of Balabanov to the combination of Akiba and Wloch would have yielded predictable results and resulted in an improved simulation of strength distribution based on material composition.
Regarding claim 9, the combination of Akiba and Wloch discloses the limitations of claim 6 and further discloses: wherein the aircraft part at least partially made of a fiber reinforced plastic having an anisotropic strength distribution is simulated by the finite element method analysis model, and an anisotropic strength is determined for each element simulating a portion made of the fiber reinforced plastic.
Akiba [0032] Meanwhile, when a material of a structural member is a composite material, such as GFRP (glass fiber reinforced plastics) or CFRP (carbon fiber reinforced plastics), it is preferable to set parameters for specifying a lamination structure of the composite material as design parameters, in addition to parameters for determining a size of the structural member made of the composite material, from a viewpoint of further minimizing weight of a structural object.
Akiba [0039]-[0042] generally disclosing patters of lamination structure (e.g., including setting a rate of the fiber reinforced layers 11 whose orientation angle is ±45° to be 40%, 50%, 60%, and 70%,, information for identifying a pattern of lamination structure, an inclined angle of the composite material 10 itself, etc., which can be set of design parameters
Akiba [0059] As described above, when plural FEM models are set as values of a design parameter, sizes and materials of structural members, such as a panel, spars, ribs, stringers, and a honeycomb structure, which compose one selected FEM model also become values of design parameters. Furthermore, when a material of a certain structural member is a composite material, a lamination structure of the composite material can also be expressed by a design parameter or design parameters as described above. As a matter of course, when a structural pattern has been fixed, only sizes, materials and the like of structural members may also be set as design parameters using a single FEM model.
Akiba [0067] Design, which prevents buckling from occurring at not more than the ultimate load, by evaluating strengths of a structural object is called non-buckling design. Nevertheless, when a material of a structural object is a composite material, the non-buckling design cannot fully take advantages of fiber reinforcing of the composite material. Thus, when a structural object is a composite material, it is desirable to perform post-buckling design which accepts buckling at not more than the ultimate load, from a viewpoint of attaining a further weight reduction of a structural object.
Akiba [0068] In the case of performing the post-buckling design of a composite material, it is necessary to estimate strengths of the composite material after buckling by performing a post-buckling analysis. Specifically, it is necessary to estimate whether strengths of a composite material after buckling have sufficient strength margins to loads applied on a structural object.
Akiba [0069] The post-buckling analysis is a nonlinear FEM analysis, of which input information is a stress distribution inside a structural object made of a composite material, for obtaining sheet thicknesses and sizes of the structural object, necessary for securing strengths required after buckling. The stress distribution inside the structural object made of a composite material, which is an input of the post-buckling analysis, can also be obtained as a result of an FEM analysis as described above.
Akiba [0071] Thus, allowable ranges of sheet thicknesses corresponding to necessary strengths after buckling, obtained as a result of a post-buckling analysis of a composite material, can be set as constraint conditions of the optimization calculation. That is, constraint conditions for securing the strengths of the composite material after the buckling can be set as the constraint conditions of the optimization calculation.
Akiba discloses strength distribution of fiber reinforced plastics, which strongly suggests, but does not explicitly state a consideration for anisotropic strength distribution. Balabanov more explicitly discloses consideration of anisotropic strength distribution when modeling composite materials (Balabanov [0041] Most composite laminates are highly anisotropic. Anisotropy can be used to control dynamic mechanical behavior in a continuum. In practice, composite laminates consist of dozens to hundreds of stacked layers or plies. It is well known that mechanical behavior of individual anisotropic layers in a composite laminate can be used to model the mechanical response of the laminate. This allows designers to tailor the elastic properties and orientation of each layer (i.e., ply) so that the mechanical response of the composite laminate will be optimized; [0081] FIG. 14 shows steps of a process for designing composite skin-stringer structures that employs optimization and finite element analysis; [0086] Referring again to FIG. 14, if the finite element analysis verifies that skin-stringer delamination will be suppressed by the candidate stringer layup design, then further analyses are performed, such as notched strength, sublaminate stability, thermal residual stresses, and interpenetration. Based on the results of these analyses, a determination is made whether the candidate stringer layup design is acceptable (step 88)..)
One of ordinary skill in the art at the time of filing would have recognized that applying the known technique of Balabanov to the combination of Akiba and Wloch would have yielded predictable results and resulted in an improved simulation of strength distribution based on material composition.
Regarding claim 15, the combination of Akiba, Wloch and Balabanov (as shown above) discloses the limitations of claim 8 and further discloses: A method of producing the aircraft part, comprising: producing the aircraft part based on the design information created by the method of designing the aircraft part according to claim 8. (Akiba [0115] Furthermore, a combination of design parameters corresponding to the optimal value of weight of the structural object can be obtained as optimal design information of the aircraft structural object. Then, an aircraft structural object can be manufactured actually in accordance with the design information of the aircraft structural object created by the above mentioned method. An aircraft structural object manufactured by such a production method becomes a structural object whose weight has been optimized. Therefore, further weight saving of the whole aircraft as a finished product can be attained)
Regarding claim 16, the combination of Akiba, Wloch and Balabanov (as shown above) discloses the limitations of claim 9 and further discloses: A method of producing the aircraft part, comprising: producing the aircraft part based on the design information created by the method of designing the aircraft part according to claim 9 (Akiba [0115] Furthermore, a combination of design parameters corresponding to the optimal value of weight of the structural object can be obtained as optimal design information of the aircraft structural object. Then, an aircraft structural object can be manufactured actually in accordance with the design information of the aircraft structural object created by the above mentioned method. An aircraft structural object manufactured by such a production method becomes a structural object whose weight has been optimized. Therefore, further weight saving of the whole aircraft as a finished product can be attained).
Conclusion
The prior art made of record and not relied upon is considered pertinent to applicant's disclosure.
Brifault (US 20220382930) discloses parameterization of a CAD model.
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/ABBY J FLYNN/ Examiner, Art Unit 3663
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