DETAILED ACTION
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 .
Priority
Acknowledgment is made of applicant’s claim for foreign priority under 35 U.S.C. 119 (a)-(d). Receipt is acknowledged of certified copies of papers required by 37 CFR 1.55.
Information Disclosure Statement
The information disclosure statement (IDS) filed 02/03/2025 has been received and considered by the examiner. The submission is in compliance with the provisions of 37 CFR 1.97.
Drawings
The drawings are objected to because the Examiner may require and is requiring descriptive text labels. Specifically, the unlabeled rectangular box(es) shown in the drawings should be provided with descriptive text labels (see Figs. 1 and 2) [MPEP 608.02(b) examiner note]. Corrected drawing sheets in compliance with 37 CFR 1.121(d) are required in reply to the Office action to avoid abandonment of the application. Any amended replacement drawing sheet should include all of the figures appearing on the immediate prior version of the sheet, even if only one figure is being amended. The figure or figure number of an amended drawing should not be labeled as “amended.” If a drawing figure is to be canceled, the appropriate figure must be removed from the replacement sheet, and where necessary, the remaining figures must be renumbered and appropriate changes made to the brief description of the several views of the drawings for consistency. Additional replacement sheets may be necessary to show the renumbering of the remaining figures. Each drawing sheet submitted after the filing date of an application must be labeled in the top margin as either “Replacement Sheet” or “New Sheet” pursuant to 37 CFR 1.121(d). If the changes are not accepted by the examiner, the applicant will be notified and informed of any required corrective action in the next Office action. The objection to the drawings will not be held in abeyance.
Claim Objections
Claims 2 and 8 are objected to because of the following informalities: the terms “the flight altitude”, “the time”, “the instruction”, “the fuel flow rate”, “the temperature”, “the rotation speed at the output”, “the output pressure”, “the torque”, “the extinction”, “the performances”, “the power demand and the power supplied”, and “the trajectory” have no previous recitation in the claims. While the scope of the claims is reasonably ascertainable, examiner recommends changing the terms to “a flight altitude”, “a time”, “an instruction”, “a fuel flow rate”, “a temperature”, “a rotation speed at an output”, “an output pressure”, “a torque”, “an extinction”, “performances”, “power demand and power supplied”, and “a trajectory”.
Appropriate correction is required.
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 1-10 are rejected under 35 U.S.C. 101 because the claimed invention is directed to an abstract idea without significantly more.
In January, 2019 (updated October 2019), the USPTO released new examination guidelines setting forth a two-step inquiry for determining whether a claim is directed to non-statutory subject matter. According to the guidelines, a claim is directed to non-statutory subject matter if:
STEP 1: the claim does not fall within one of the four statutory categories of invention (process, machine, manufacture or composition of matter), or
STEP 2: the claim recites a judicial exception, e.g. an abstract idea, without reciting additional elements that amount to significantly more than the judicial exception, as determined using the following analysis:
STEP 2A (PRONG 1): Does the claim recite an abstract idea, law of nature, or natural phenomenon?
STEP 2A (PRONG 2): Does the claim recite additional elements that integrate the judicial exception into a practical application?
STEP 2B: Does the claim recite additional elements that amount to significantly more than the judicial exception?
Using the two-step inquiry, it is clear that claim 1 is directed toward non-statutory subject matter, as shown below:
STEP 1: Does claim 1 fall within one of the statutory categories? Yes. The claim is directed toward a method including at least one step.
STEP 2A (PRONG 1): Is the claim directed to a law of nature, a natural phenomenon or an abstract idea? Yes, the claim is directed to an abstract idea.
With regard to STEP 2A (PRONG 1), the guidelines provide three groupings of subject matter that are considered abstract ideas:
Mathematical concepts – mathematical relationships, mathematical formulas or equations, mathematical calculations;
Certain methods of organizing human activity – fundamental economic principles or practices (including hedging, insurance, mitigating risk); commercial or legal interactions (including agreements in the form of contracts; legal obligations; advertising, marketing or sales activities or behaviors; business relations); managing personal behavior or relationships or interactions between people (including social activities, teaching, and following rules or instructions); and
Mental processes – concepts that are practicably performed in the human mind (including an observation, evaluation, judgment, opinion).
Claim 1. Propulsion-assistance method by detecting a failure in a turboshaft engine of an aircraft operating in a nominal mode, the aircraft comprising a deactivated assistance engine, the method comprising a step of comparing operating parameters of the turboshaft engine with equivalent parameters of a model representing a healthy turboshaft engine, a step of detecting a failure in the turboshaft engine by detecting an anomaly in at least one operating parameter of the turboshaft engine, a step of selecting an activation mode of the assistance engine according to the operating parameters of the turboshaft engine and/or flight parameters of the aircraft, and a step of activating the assistance engine with the activation mode selected.
The method in claim 1, specifically the limitations emphasized above, is a mental process that can be practicably performed in the human mind and, therefore, recites an abstract idea. The claimed method merely consists of comparing operating parameters, detecting a failure, selecting an activation mode, and activating the assistance engine. This is equivalent to a person mentally evaluating operating parameters, determining whether a failure has occurred, and determining an activation mode. Thus, claim 1 recite a mental process.
STEP 2A (PRONG 2): Does the claim recite additional elements that integrate the judicial exception into a practical application? No, the claim does not recite additional elements that integrate the judicial exception into a practical application.
With regard to STEP 2A (prong 2), whether the claim recites additional elements that integrate the judicial exception into a practical application, the guidelines provide the following exemplary considerations that are indicative that an additional element (or combination of elements) may have integrated the judicial exception into a practical application:
an additional element reflects an improvement in the functioning of a computer, or an improvement to other technology or technical field;
an additional element that applies or uses a judicial exception to effect a particular treatment or prophylaxis for a disease or medical condition;
an additional element implements a judicial exception with, or uses a judicial exception in conjunction with, a particular machine or manufacture that is integral to the claim;
an additional element effects a transformation or reduction of a particular article to a different state or thing; and
an additional element applies or uses the judicial exception in some other meaningful way beyond generally linking the use of the judicial exception to a particular technological environment, such that the claim as a whole is more than a drafting effort designed to monopolize the exception.
While the guidelines further state that the exemplary considerations are not an exhaustive list and that there may be other examples of integrating the exception into a practical application, the guidelines also list examples in which a judicial exception has not been integrated into a practical application:
an additional element merely recites the words “apply it” (or an equivalent) with the judicial exception, or merely includes instructions to implement an abstract idea on a computer, or merely uses a computer as a tool to perform an abstract idea;
an additional element adds insignificant extra-solution activity to the judicial exception; and
an additional element does no more than generally link the use of a judicial exception to a particular technological environment or field of use.
Claim 1. Propulsion-assistance method by detecting a failure in a turboshaft engine of an aircraft operating in a nominal mode, the aircraft comprising a deactivated assistance engine, the method comprising a step of comparing operating parameters of the turboshaft engine with equivalent parameters of a model representing a healthy turboshaft engine, a step of detecting a failure in the turboshaft engine by detecting an anomaly in at least one operating parameter of the turboshaft engine, a step of selecting an activation mode of the assistance engine according to the operating parameters of the turboshaft engine and/or flight parameters of the aircraft, and a step of activating the assistance engine with the activation mode selected.
Claim 1 does not recite any of the exemplary considerations that are indicative of an abstract idea having been integrated into a practical application. The additional elements are underlined in the claim above. Specifically, the activating step is recited at a high level of generality (i.e. as a activation of the assistance engine which under a BRI, could be a person starting or activating the assistance engine as the specification in the last paragraph at page 10 states that “in one embodiment, the method is implemented continuously by a computer” which is open ended and thus leaves open that the method may not always be implemented by a computer) and amount to mere post solution actions, which is a form of insignificant extra solution activity. As such, claim 1 is not integrated into practical application.
STEP 2B: Does the claim recite additional elements that amount to significantly more than the judicial exception? No, the claims do not recite additional elements that amount to significantly more than the judicial exception.
With regard to STEP 2B, whether the claims recite additional elements that provide significantly more than the recited judicial exception, the guidelines specify that the pre-guideline procedure is still in effect. Specifically, that examiners should continue to consider whether an additional element or combination of elements:
adds a specific limitation or combination of limitations that are not well-understood, routine, conventional activity in the field, which is indicative that an inventive concept may be present; or
simply appends well-understood, routine, conventional activities previously known to the industry, specified at a high level of generality, to the judicial exception, which is indicative that an inventive concept may not be present.
Regarding Step 2B of the 2019 PEG, independent claim 1 does not include additional elements (considered both individually and as an ordered combination) that are sufficient to amount to significantly more than the judicial exception for the same reasons discussed above with respect to determining that the claims do not integrate the abstract idea into a practical application.
Thus, since claim 1 is: (a) directed toward an abstract idea, (b) does not recite additional elements that integrate the judicial exception into a practical application, and (c) does not recite additional elements that amount to significantly more than the judicial exception, it is clear that claim 1 is directed towards non-statutory subject matter.
Dependent claims 2-9 further limit the abstract idea without integrating the abstract idea into practical application or adding significantly more, such as the limitations in claim 2 that amount to insignificant extra solution activity using a similar analysis applied to claim 1 above.
As such, claims 1-10 are rejected under 35 USC 101 as being drawn to an abstract idea without significantly more, and thus are ineligible.
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.
This application currently names joint inventors. In considering patentability of the claims the examiner presumes that the subject matter of the various claims was commonly owned as of the effective filing date of the claimed invention(s) absent any evidence to the contrary. Applicant is advised of the obligation under 37 CFR 1.56 to point out the inventor and effective filing dates of each claim that was not commonly owned as of the effective filing date of the later invention in order for the examiner to consider the applicability of 35 U.S.C. 102(b)(2)(C) for any potential 35 U.S.C. 102(a)(2) prior art against the later invention.
Claim(s) 1-4, 6, 8, 10 is/are rejected under 35 U.S.C. 103 as being unpatentable over Vallart (US 20220024568 A1) in view of Vhora (US 20050222747 A1).
Regarding Claim 1, Vallart teaches Propulsion-assistance method by detecting a failure in a turboshaft engine of an aircraft operating in a nominal mode (See at least paragraph [0009], “An exit from standby mode is described as “normal” when a change in flight situation requires the activation of the turboshaft engine that is on standby, for example when the helicopter is going to transition from a cruise flight situation to a landing phase. A normal exit from standby of this kind occurs over a period of time between 10 seconds and 1 minute. An exit from standby mode is described as “rapid” when a failure or deficit of power in the active engine occurs or when the flight conditions suddenly become difficult. An emergency exit from standby of this kind occurs over a period of less than 10 seconds” and paragraph [0021], “A hybrid turboshaft engine within the meaning of the disclosure is a turboshaft engine configured to be capable of being put, on demand and deliberately, in at least one predetermined standby mode, from which it can exit in a normal or rapid (also referred to as emergency) manner. A turboshaft engine can be in standby mode only during a stable flight of the helicopter, i.e., when no turboshaft engine of the helicopter has failed, during a cruise flight situation, when it is progressing in normal conditions. The exit from standby mode consists in changing the turboshaft engine into a gas generator acceleration mode by means of driving in a manner that is compatible with the exit mode required by the conditions (normal standby-exiting mode or rapid standby-exiting mode, also referred to as emergency exit).”), the aircraft comprising a deactivated assistance engine (See at least paragraph [0018], “at least one turboshaft engine among the turboshaft engines, referred to as a hybrid turboshaft engine, capable of operating in at least one standby mode during a stable flight of the helicopter, the other turboshaft engines operating alone during this stable flight” and paragraph [0021], “A hybrid turboshaft engine within the meaning of the disclosure is a turboshaft engine configured to be capable of being put, on demand and deliberately, in at least one predetermined standby mode, from which it can exit in a normal or rapid (also referred to as emergency) manner. A turboshaft engine can be in standby mode only during a stable flight of the helicopter, i.e., when no turboshaft engine of the helicopter has failed, during a cruise flight situation, when it is progressing in normal conditions. The exit from standby mode consists in changing the turboshaft engine into a gas generator acceleration mode by means of driving in a manner that is compatible with the exit mode required by the conditions (normal standby-exiting mode or rapid standby-exiting mode, also referred to as emergency exit).”), a step of detecting a failure in the turboshaft engine by detecting an anomaly in at least one operating parameter of the turboshaft engine (See at least paragraph [0009], “An exit from standby mode is described as “normal” when a change in flight situation requires the activation of the turboshaft engine that is on standby, for example when the helicopter is going to transition from a cruise flight situation to a landing phase. A normal exit from standby of this kind occurs over a period of time between 10 seconds and 1 minute. An exit from standby mode is described as “rapid” when a failure or deficit of power in the active engine occurs or when the flight conditions suddenly become difficult. An emergency exit from standby of this kind occurs over a period of less than 10 seconds.”), a step of selecting an activation mode of the assistance engine according to the operating parameters of the turboshaft engine and/or flight parameters of the aircraft (See at least paragraph [0009], “An exit from standby mode is described as “normal” when a change in flight situation requires the activation of the turboshaft engine that is on standby, for example when the helicopter is going to transition from a cruise flight situation to a landing phase. A normal exit from standby of this kind occurs over a period of time between 10 seconds and 1 minute. An exit from standby mode is described as “rapid” when a failure or deficit of power in the active engine occurs or when the flight conditions suddenly become difficult. An emergency exit from standby of this kind occurs over a period of less than 10 seconds” and paragraph [0021], “A hybrid turboshaft engine within the meaning of the disclosure is a turboshaft engine configured to be capable of being put, on demand and deliberately, in at least one predetermined standby mode, from which it can exit in a normal or rapid (also referred to as emergency) manner. A turboshaft engine can be in standby mode only during a stable flight of the helicopter, i.e., when no turboshaft engine of the helicopter has failed, during a cruise flight situation, when it is progressing in normal conditions. The exit from standby mode consists in changing the turboshaft engine into a gas generator acceleration mode by means of driving in a manner that is compatible with the exit mode required by the conditions (normal standby-exiting mode or rapid standby-exiting mode, also referred to as emergency exit).”), and a step of activating the assistance engine with the activation mode selected (See at least paragraph [0009], “An exit from standby mode is described as “normal” when a change in flight situation requires the activation of the turboshaft engine that is on standby, for example when the helicopter is going to transition from a cruise flight situation to a landing phase. A normal exit from standby of this kind occurs over a period of time between 10 seconds and 1 minute. An exit from standby mode is described as “rapid” when a failure or deficit of power in the active engine occurs or when the flight conditions suddenly become difficult. An emergency exit from standby of this kind occurs over a period of less than 10 seconds” and paragraph [0021], “A hybrid turboshaft engine within the meaning of the disclosure is a turboshaft engine configured to be capable of being put, on demand and deliberately, in at least one predetermined standby mode, from which it can exit in a normal or rapid (also referred to as emergency) manner. A turboshaft engine can be in standby mode only during a stable flight of the helicopter, i.e., when no turboshaft engine of the helicopter has failed, during a cruise flight situation, when it is progressing in normal conditions. The exit from standby mode consists in changing the turboshaft engine into a gas generator acceleration mode by means of driving in a manner that is compatible with the exit mode required by the conditions (normal standby-exiting mode or rapid standby-exiting mode, also referred to as emergency exit).”).
Vallart does not explicitly disclose, however, Vhora, in the same field of endeavor, teaches the method comprising a step of comparing operating parameters of the turboshaft engine with equivalent parameters of a model representing a healthy turboshaft engine (See at least paragraph [0029], “The gas path model 15 is shown in FIG. 2 in detail, where external inputs 21, including customer bleeds, shaft power load, and starter torque are sent from sensors (in external inputs 21) to a CMEM model 23, which is a Component Map-based Engine Model and it is a nonlinear model. The incoming engine data is stored and processed. Ambient operating inputs 25 including, for example, temperature, pressure, Mach number and speed. Engine control unit inputs 27 are input into CMEM 23, including fuel commands and surge bleed valve control elements. Residuals are calculated by comparing the actual data inputs (on a continuous or steady state basis) with model predicted values that model CMEM 23 has determined for a set of data that represents the operating conditions of the engine being monitored. Model outputs 29 are computed for all the operating conditions, including shaft speeds, all temperatures, all pressures and all air flows” and paragraph [0032], “Various engine component faults are modeled by modifying one or more component characteristics in the model. For example, the high pressure compressor deteriorates in performance due to erosion of its rotor blades. This erosion occurs faster if an engine operates in areas with high air borne dust particles. This degradation manifests in lost performance in terms of reduction in efficiency, air flow and pressure rise at given conditions. This type of degradation or fault is modeled by reducing the efficiency of the compressor in the model, along with reducing its air flow and pressure rise characteristics. Thus the fault models represent the actual physics of the fault, based on fundamental knowledge of the components. These models are also validated by matching them with test data from actually degraded engines. All these parameters are varied in fixed proportion,, such that each fault manifests itself by unique variation in output parameters, compared to normal or healthy models. Typical faults that have been modeled are HP turbine deterioration, bleed band fault, HP compressor deterioration, fan rotor deterioration, LP compressor deterioration and air/oil cooler clogging in the lube oil system.”).
Thus, it would have been obvious to one of ordinary skill in the art before the effective filing date to combine the invention of Vallart with the teachings of Vhora such that the helicopter propulsion system of Vallart is further configured to compare operating parameters of the turboshaft engine with equivalent parameters of a model representing a healthy turboshaft engine, as taught by Vhora (See paragraph [0029], [0032].), with a reasonable expectation of success. The motivation for doing so would be to detect existing engine performance-related faults using a nonlinear engine model in conjunction with operational engine data, as taught by Vhora (See paragraph [0016].).
Regarding Claim 2, Vallart and Vhora teach Method according to claim 1, as set forth in the obviousness rejection. Vallart teaches wherein the flight parameters of the aircraft comprise the flight altitude of the aircraft (See at least paragraph [0006], “The turboshaft engines are also oversized in order to be able to ensure flight over the entire flight range specified by the aircraft manufacturer and in particular flight at high altitudes and in hot weather. These flight points, which are highly demanding, particularly when the weight of the helicopter is close to its maximum take-off weight, are encountered only in certain circumstances of use.”), the activation mode of the assistance engine being selected according to the possibility of maintaining the flight altitude of the aircraft over the time necessary for activating the assistance engine (See at least paragraph [0005], “Furthermore, the turboshaft engines of a helicopter are designed so as to be oversized in order to be able to keep the helicopter in flight in the event that one of the engines fails. This flight situation arises after the loss of an engine and results in the fact that each operating engine supplies a power level much beyond its nominal power so that the helicopter can cope with a dangerous situation, and then continue its flight”, paragraph [0006], “The turboshaft engines are also oversized in order to be able to ensure flight over the entire flight range specified by the aircraft manufacturer and in particular flight at high altitudes and in hot weather. These flight points, which are highly demanding, particularly when the weight of the helicopter is close to its maximum take-off weight, are encountered only in certain circumstances of use”, paragraph [0009], “An exit from standby mode is described as “normal” when a change in flight situation requires the activation of the turboshaft engine that is on standby, for example when the helicopter is going to transition from a cruise flight situation to a landing phase. A normal exit from standby of this kind occurs over a period of time between 10 seconds and 1 minute. An exit from standby mode is described as “rapid” when a failure or deficit of power in the active engine occurs or when the flight conditions suddenly become difficult. An emergency exit from standby of this kind occurs over a period of less than 10 seconds”, and paragraph [0021], “A hybrid turboshaft engine within the meaning of the disclosure is a turboshaft engine configured to be capable of being put, on demand and deliberately, in at least one predetermined standby mode, from which it can exit in a normal or rapid (also referred to as emergency) manner. A turboshaft engine can be in standby mode only during a stable flight of the helicopter, i.e., when no turboshaft engine of the helicopter has failed, during a cruise flight situation, when it is progressing in normal conditions. The exit from standby mode consists in changing the turboshaft engine into a gas generator acceleration mode by means of driving in a manner that is compatible with the exit mode required by the conditions (normal standby-exiting mode or rapid standby-exiting mode, also referred to as emergency exit).”).
Regarding Claim 3, Vallart and Vhora teach Method according to claim 1, as set forth in the obviousness rejection. Vallart does not explicitly disclose, however, Vhora, in the same field of endeavor, teaches wherein the step of comparing operating parameters of the turboshaft engine with the equivalent parameters of a model representing a healthy turboshaft engine is implemented by a computer of the turboshaft engine storing the model representing a healthy turboshaft engine (See at least paragraph [0017], “The residuals are statistically analyzed to estimate bounds of uncertainties as indicative of sensor noise. Then the incoming residuals are compared from engine data against bounds, such that a fault is detected when a threshold is exceeded. Detection of a fault activates a computer to calculate fault residuals for each fault model, using a model matching technique. At this point, the fault model which brings the residuals back to normal bounds is the diagnosed fault”, paragraph [0029], “The gas path model 15 is shown in FIG. 2 in detail, where external inputs 21, including customer bleeds, shaft power load, and starter torque are sent from sensors (in external inputs 21) to a CMEM model 23, which is a Component Map-based Engine Model and it is a nonlinear model. The incoming engine data is stored and processed. Ambient operating inputs 25 including, for example, temperature, pressure, Mach number and speed. Engine control unit inputs 27 are input into CMEM 23, including fuel commands and surge bleed valve control elements. Residuals are calculated by comparing the actual data inputs (on a continuous or steady state basis) with model predicted values that model CMEM 23 has determined for a set of data that represents the operating conditions of the engine being monitored. Model outputs 29 are computed for all the operating conditions, including shaft speeds, all temperatures, all pressures and all air flows”, paragraph [0032], “Various engine component faults are modeled by modifying one or more component characteristics in the model. For example, the high pressure compressor deteriorates in performance due to erosion of its rotor blades. This erosion occurs faster if an engine operates in areas with high air borne dust particles. This degradation manifests in lost performance in terms of reduction in efficiency, air flow and pressure rise at given conditions. This type of degradation or fault is modeled by reducing the efficiency of the compressor in the model, along with reducing its air flow and pressure rise characteristics. Thus the fault models represent the actual physics of the fault, based on fundamental knowledge of the components. These models are also validated by matching them with test data from actually degraded engines. All these parameters are varied in fixed proportion,, such that each fault manifests itself by unique variation in output parameters, compared to normal or healthy models. Typical faults that have been modeled are HP turbine deterioration, bleed band fault, HP compressor deterioration, fan rotor deterioration, LP compressor deterioration and air/oil cooler clogging in the lube oil system.”).
Thus, it would have been obvious to one of ordinary skill in the art before the effective filing date to combine the invention of Vallart with the teachings of Vhora such that the helicopter propulsion system of Vallart is further configured to compare operating parameters of the turboshaft engine with equivalent parameters of a model representing a healthy turboshaft engine which is implemented by a computer of the turboshaft engine storing the model representing a healthy turboshaft engine, as taught by Vhora (See paragraph [0017], [0029], [0032].), with a reasonable expectation of success. The motivation for doing so would be to detect existing engine performance-related faults using a nonlinear engine model in conjunction with operational engine data, as taught by Vhora (See paragraph [0016].).
Regarding Claim 4, Vallart and Vhora teach Method according to claim 1, as set forth in the obviousness rejection. Vallart does not explicitly disclose, however, Vhora, in the same field of endeavor, teaches wherein the step of detecting a failure in the turboshaft engine comprises measuring a significant difference between an operating parameter of the turboshaft engine and the equivalent parameter of the model representing a healthy turboshaft engine (See at least paragraph [0017], “The residuals are statistically analyzed to estimate bounds of uncertainties as indicative of sensor noise. Then the incoming residuals are compared from engine data against bounds, such that a fault is detected when a threshold is exceeded. Detection of a fault activates a computer to calculate fault residuals for each fault model, using a model matching technique. At this point, the fault model which brings the residuals back to normal bounds is the diagnosed fault”, paragraph [0029], “The gas path model 15 is shown in FIG. 2 in detail, where external inputs 21, including customer bleeds, shaft power load, and starter torque are sent from sensors (in external inputs 21) to a CMEM model 23, which is a Component Map-based Engine Model and it is a nonlinear model. The incoming engine data is stored and processed. Ambient operating inputs 25 including, for example, temperature, pressure, Mach number and speed. Engine control unit inputs 27 are input into CMEM 23, including fuel commands and surge bleed valve control elements. Residuals are calculated by comparing the actual data inputs (on a continuous or steady state basis) with model predicted values that model CMEM 23 has determined for a set of data that represents the operating conditions of the engine being monitored. Model outputs 29 are computed for all the operating conditions, including shaft speeds, all temperatures, all pressures and all air flows”, paragraph [0032], “Various engine component faults are modeled by modifying one or more component characteristics in the model. For example, the high pressure compressor deteriorates in performance due to erosion of its rotor blades. This erosion occurs faster if an engine operates in areas with high air borne dust particles. This degradation manifests in lost performance in terms of reduction in efficiency, air flow and pressure rise at given conditions. This type of degradation or fault is modeled by reducing the efficiency of the compressor in the model, along with reducing its air flow and pressure rise characteristics. Thus the fault models represent the actual physics of the fault, based on fundamental knowledge of the components. These models are also validated by matching them with test data from actually degraded engines. All these parameters are varied in fixed proportion,, such that each fault manifests itself by unique variation in output parameters, compared to normal or healthy models. Typical faults that have been modeled are HP turbine deterioration, bleed band fault, HP compressor deterioration, fan rotor deterioration, LP compressor deterioration and air/oil cooler clogging in the lube oil system.”).
Thus, it would have been obvious to one of ordinary skill in the art before the effective filing date to combine the invention of Vallart with the teachings of Vhora such that the helicopter propulsion system of Vallart is further configured to compare operating parameters of the turboshaft engine with equivalent parameters of a model representing a healthy turboshaft engine and detect a failure in the turboshaft engine comprising measuring a significant difference between an operating parameter of the turboshaft engine and the equivalent parameter of the model representing a healthy turboshaft engine, as taught by Vhora (See paragraph [0017], [0029], [0032].), with a reasonable expectation of success. The motivation for doing so would be to detect existing engine performance-related faults using a nonlinear engine model in conjunction with operational engine data, as taught by Vhora (See paragraph [0016].).
Regarding Claim 6, Vallart and Vhora teach Method according to claim 1, as set forth in the obviousness rejection. Vallart teaches wherein the assistance engine is a second turboshaft engine (See at least paragraph [0017]-[0018], “In order to achieve this, the disclosure relates to an architecture of a propulsion system of a multi-engine helicopter comprising turboshaft engines connected to a power transmission gearbox, characterized in that it comprises: at least one turboshaft engine among the turboshaft engines, referred to as a hybrid turboshaft engine, capable of operating in at least one standby mode during a stable flight of the helicopter, the other turboshaft engines operating alone during this stable flight.”).
Regarding Claim 8, Vallart and Vhora teach Method according to claim 1, as set forth in the obviousness rejection. Vallart does not explicitly disclose, however, Vhora, in the same field of endeavor, teaches wherein the operating parameters of the turboshaft engine or the detectable failures comprise the fuel flow rate, and/or the temperature and/or the rotation speed at the output of a high-pressure turbine of the turboshaft engine, and/or the output pressure of a compressor of the turboshaft engine, and/or the torque (See at least paragraph [0029], “The gas path model 15 is shown in FIG. 2 in detail, where external inputs 21, including customer bleeds, shaft power load, and starter torque are sent from sensors (in external inputs 21) to a CMEM model 23, which is a Component Map-based Engine Model and it is a nonlinear model. The incoming engine data is stored and processed. Ambient operating inputs 25 including, for example, temperature, pressure, Mach number and speed. Engine control unit inputs 27 are input into CMEM 23, including fuel commands and surge bleed valve control elements. Residuals are calculated by comparing the actual data inputs (on a continuous or steady state basis) with model predicted values that model CMEM 23 has determined for a set of data that represents the operating conditions of the engine being monitored. Model outputs 29 are computed for all the operating conditions, including shaft speeds, all temperatures, all pressures and all air flows.”), and/or the extinction of a combustion chamber of the turboshaft engine, and/or a leak on an oil circuit, and/or a leak on a fuel circuit, and/or a detection of chips, and/or an inability to regulate the performances of the turboshaft engine, and/or inconsistency between the power demand and the power supplied, and/or a potential impact on the trajectory of the aircraft.
Thus, it would have been obvious to one of ordinary skill in the art before the effective filing date to combine the invention of Vallart with the teachings of Vhora such that the helicopter propulsion system of Vallart is further configured to compare operating parameters of the turboshaft engine with equivalent parameters of a model representing a healthy turboshaft engine and the operating parameters of the turboshaft engine or the detectable failures comprise the fuel flow rate, the temperature, the rotation speed at the output of a high-pressure turbine of the turboshaft engine, the output pressure of a compressor of the turboshaft engine, and the torque, as taught by Vhora (See paragraph [0029], [0032].), with a reasonable expectation of success. The motivation for doing so would be to detect existing engine performance-related faults using a nonlinear engine model in conjunction with operational engine data, as taught by Vhora (See paragraph [0016].).
Regarding Claim 10, Vallart and Vhora teach Method according to claim 1, as set forth in the obviousness rejection. Vallart teaches wherein the deactivated assistance engine corresponds to an operating mode of the assistance engine equivalent to a standby mode (See at least paragraph [0018], “at least one turboshaft engine among the turboshaft engines, referred to as a hybrid turboshaft engine, capable of operating in at least one standby mode during a stable flight of the helicopter, the other turboshaft engines operating alone during this stable flight” and paragraph [0021], “A hybrid turboshaft engine within the meaning of the disclosure is a turboshaft engine configured to be capable of being put, on demand and deliberately, in at least one predetermined standby mode, from which it can exit in a normal or rapid (also referred to as emergency) manner. A turboshaft engine can be in standby mode only during a stable flight of the helicopter, i.e., when no turboshaft engine of the helicopter has failed, during a cruise flight situation, when it is progressing in normal conditions. The exit from standby mode consists in changing the turboshaft engine into a gas generator acceleration mode by means of driving in a manner that is compatible with the exit mode required by the conditions (normal standby-exiting mode or rapid standby-exiting mode, also referred to as emergency exit).”) or to a mode intermediate between a standby mode and a nominal operating mode.
Claim(s) 5 is/are rejected under 35 U.S.C. 103 as being unpatentable over Vallart (US 20220024568 A1) in view of Vhora (US 20050222747 A1) and MARCONI (US 20170122221 A1).
Regarding Claim 5, Vallart and Vhora teach Method according to claim 1, as set forth in the obviousness rejection. Vallart teaches wherein the activation mode of the assistance engine is selected between a normal activation mode having a first activation duration (See at least paragraph [0019], “at least two systems for controlling each hybrid turboshaft engine, referred to as reactivation systems, each system comprising an electric machine connected to the hybrid turboshaft engine and designed to be capable of rotating the hybrid turboshaft engine, and at least one source of electrical power for the electric machine, each reactivation system being configured such that it can drive the turboshaft engine in at least one operating mode among a plurality of predetermined modes” and paragraph [0024], “a mode, referred to as the normal reactivation mode, in which the turboshaft engine is rotated from the standby mode up to a speed in the range of between 80 and 105% of the nominal speed of the gas generator of the turboshaft engine in a period in the range of between 10 seconds and 60 seconds.”), and a rapid activation mode having a third activation duration (See at least paragraph [0023], “a mode, referred to as the rapid reactivation mode, in which the turboshaft engine is rotated from the standby mode up to a speed in the range of between 80 and 105% of the nominal speed of the gas generator of the turboshaft engine in a period of less than 10 seconds.”), the first activation duration being longer than…the third activation duration (See at least paragraph [0023], “a mode, referred to as the rapid reactivation mode, in which the turboshaft engine is rotated from the standby mode up to a speed in the range of between 80 and 105% of the nominal speed of the gas generator of the turboshaft engine in a period of less than 10 seconds” and paragraph [0024], “a mode, referred to as the normal reactivation mode, in which the turboshaft engine is rotated from the standby mode up to a speed in the range of between 80 and 105% of the nominal speed of the gas generator of the turboshaft engine in a period in the range of between 10 seconds and 60 seconds.”), the normal activation mode being favoured over…the rapid activation mode (See at least paragraph [0023], “a mode, referred to as the rapid reactivation mode, in which the turboshaft engine is rotated from the standby mode up to a speed in the range of between 80 and 105% of the nominal speed of the gas generator of the turboshaft engine in a period of less than 10 seconds”, paragraph [0024], “a mode, referred to as the normal reactivation mode, in which the turboshaft engine is rotated from the standby mode up to a speed in the range of between 80 and 105% of the nominal speed of the gas generator of the turboshaft engine in a period in the range of between 10 seconds and 60 seconds”, paragraph [0035], “According to this variant, the first and the second reactivation systems are both compatible with a normal reactivation of the turboshaft engine. They can therefore be called upon alternately at each start-up in order to check their availability”, and paragraph [0037], “In the event that the first system is unavailable, the second system is called upon for a normal reactivation of the hybrid turboshaft engine.” Normal reactivation is the routine operating mode because both reactivation systems are compatible with normal reactivation and are called upon alternately at each start-up, whereas the first reactivation system is maintained in readiness for rapid reactivation.).
Vallart and Vhora do not explicitly disclose, however, MARCONI, in the same field of endeavor, teaches an accelerated activation mode having a second activation duration (See at least paragraph [0009]-[0011], “The applicants have proposed in particular the following two standby modes: a standby mode referred to as normal super-idling, in which the combustion chamber is ignited and the shaft of the gas generator rotates at a speed of between 20 and 60% of the nominal speed, a standby mode referred to as assisted super-idling, in which the combustion chamber is ignited and the shaft of the gas generator rotates, with mechanical assistance, at a speed of between 20 and 60% of the nominal speed.”), …the second activation duration… (See at least paragraph [0009]-[0011], “The applicants have proposed in particular the following two standby modes: a standby mode referred to as normal super-idling, in which the combustion chamber is ignited and the shaft of the gas generator rotates at a speed of between 20 and 60% of the nominal speed, a standby mode referred to as assisted super-idling, in which the combustion chamber is ignited and the shaft of the gas generator rotates, with mechanical assistance, at a speed of between 20 and 60% of the nominal speed.”), …the accelerated activation mode also being favoured over the rapid activation mode (See at least paragraph [0009]-[0013], “The applicants have proposed in particular the following two standby modes: a standby mode referred to as normal super-idling, in which the combustion chamber is ignited and the shaft of the gas generator rotates at a speed of between 20 and 60% of the nominal speed, a standby mode referred to as assisted super-idling, in which the combustion chamber is ignited and the shaft of the gas generator rotates, with mechanical assistance, at a speed of between 20 and 60% of the nominal speed. A disadvantage of the normal super-idling mode is the operating temperatures, which become increasingly high as attempts are made to reach ever lower idling. The assisted super-idling mode makes it possible to remedy this problem of operating temperature. However, this requires the use of an electrical or pneumatic drive machine and of a corresponding coupling.” The assisted super-idling mode remedies the operating-temperature disadvantage of the normal super-idling mode.).
Thus, it would have been obvious to one of ordinary skill in the art before the effective filing date to combine the invention of Vallart with the teachings of Vhora and Marconi such that the helicopter propulsion system of Vallart is further configured to compare operating parameters of the turboshaft engine with equivalent parameters of a model representing a healthy turboshaft engine, as taught by Vhora (See paragraph [0029], [0032].), and to utilize an accelerated activation mode having a second activation duration, the accelerated activation mode also being favoured over the rapid activation mode, as taught by Marconi (See paragraphs [0009]-[0013].), with a reasonable expectation of success. The motivation for doing so would be to detect existing engine performance-related faults using a nonlinear engine model in conjunction with operational engine data, as taught by Vhora (See paragraph [0016].), and to remedy the operating-temperature disadvantage of the normal super-idling mode while maintaining the engine in a standby mode, as taught by Marconi (See paragraph [0012], [0013].).
Claim(s) 7 and 9 is/are rejected under 35 U.S.C. 103 as being unpatentable over Vallart (US 20220024568 A1) in view of Vhora (US 20050222747 A1) and MARCONI (US 20170122221 A1) and BEDRINE (US 20230193779 A1).
Regarding Claim 7, Vallart, Vhora, and MARCONI teach Method according to claim 5, as set forth in the obviousness rejection. Vallart, Vhora, and MARCONI do not explicitly disclose, however, BEDRINE, in the same field of endeavor, teaches wherein the normal activation mode comprises a step of thermal stabilisation at a predefined temperature of the assistance engine (See at least paragraph [0021], “In a second aspect of the method for quickly stopping the helicopter rotor, the rotation of the extinguished gas generator by said at least one electrical machine can be maintained as long as at least one measured temperature of the gas generator is greater than a first threshold and/or the absolute value of a measured temperature gradient is greater than a second threshold”, paragraph [0022], “In one aspect, the rotation of the extinguished gas generator can thus be maintained for a predefined ventilation period, but shut off before the expiration of this ventilation period as soon as the measured temperature of the gas generator is less than the first threshold”, paragraph [0049], “In a second step 110 of the method, an electronic control unit determines whether the thermal stabilization phase of the gas generator 12 of the turbomachine 4 has already taken place. If it has already taken place, in a step 115 the control unit transmits a signal to control the extinction of the gas generator 12, following which the pilot commands the stop of the main rotor 2 by the application of a brake on the shaft 6 of the main rotor 2”, and paragraph [0050], “If, on the other hand, no thermal stabilization phase is detected, the control unit commands the extinction of the combustion chamber of the gas generator 12 of the turbomachine 4 in a step 120, then the braking of the main rotor 2 in a step 130, and the ventilation of the gas generator 12 of the turbomachine in a step 140. The ventilation of the gas generator 12 is accomplished by the rotation of the gas generator 12 by an electrical machine which is powered by the electrical network of the helicopter. The electrical network of the helicopter can be coupled to one or more battery(ies), and possibly to at least one auxiliary power unit.” The thermal stabilization phase is controlled using a measured temperature of the gas generator relative to a first threshold, which corresponds to a predefined temperature criterion for the engine.) and is favoured when no failure of the turboshaft engine is detected and an onboard computer and/or a pilot of the aircraft gives the instruction for this to the assistance engine (See at least paragraph [0048], “In a first step 100 of the method, at the conclusion of a landing phase of the helicopter 1, the pilot issues a command to stop the engine in order to quickly accomplish a given mission such as for example disembarking or embarking passengers” and paragraph [0049], “In a second step 110 of the method, an electronic control unit determines whether the thermal stabilization phase of the gas generator 12 of the turbomachine 4 has already taken place. If it has already taken place, in a step 115 the control unit transmits a signal to control the extinction of the gas generator 12, following which the pilot commands the stop of the main rotor 2 by the application of a brake on the shaft 6 of the main rotor 2.”).
Thus, it would have been obvious to one of ordinary skill in the art before the effective filing date to combine the invention of Vallart with the teachings of Vhora, Marconi, and BEDRINE such that the helicopter propulsion system of Vallart is further configured to compare operating parameters of the turboshaft engine with equivalent parameters of a model representing a healthy turboshaft engine, as taught by Vhora (See paragraph [0029], [0032].), to utilize an accelerated activation mode having a second activation duration, the accelerated activation mode also being favoured over the rapid activation mode, as taught by Marconi (See paragraphs [0009]-[0013].), and the normal activation mode comprises a step of thermal stabilisation at a predefined temperature of the assistance engine and is favoured when a pilot of the aircraft gives the instruction for this to the assistance engine, as taught by BEDRINE (See paragraph [0021], [0022], [0048]-[0050].), with a reasonable expectation of success. The motivation for doing so would be to detect existing engine performance-related faults using a nonlinear engine model in conjunction with operational engine data, as taught by Vhora (See paragraph [0016].), to remedy the operating-temperature disadvantage of the normal super-idling mode while maintaining the engine in a standby mode, as taught by Marconi (See paragraph [0012], [0013].), and to protect the engine against mechanical stresses associated with strong variations in power demand during thermal stabliziation, as taught by BEDRINE (See paragraph [0004].).
Regarding Claim 9, Vallart, Vhora, and MARCONI teach Method according to claim 5, as set forth in the obviousness rejection. Vallart teaches wherein the assistance engine is a second turboshaft engine (See at least paragraph [0017]-[0018], “In order to achieve this, the disclosure relates to an architecture of a propulsion system of a multi-engine helicopter comprising turboshaft engines connected to a power transmission gearbox, characterized in that it comprises: at least one turboshaft engine among the turboshaft engines, referred to as a hybrid turboshaft engine, capable of operating in at least one standby mode during a stable flight of the helicopter, the other turboshaft engines operating alone during this stable flight.”), and wherein the normal activation mode…and the rapid activation mode comprise the assistance of the second turboshaft engine by an electric machine of the aircraft (See at least paragraph [0019], “at least two systems for controlling each hybrid turboshaft engine, referred to as reactivation systems, each system comprising an electric machine connected to the hybrid turboshaft engine and designed to be capable of rotating the hybrid turboshaft engine, and at least one source of electrical power for the electric machine, each reactivation system being configured such that it can drive the turboshaft engine in at least one operating mode among a plurality of predetermined modes”, paragraph [0023], “a mode, referred to as the rapid reactivation mode, in which the turboshaft engine is rotated from the standby mode up to a speed in the range of between 80 and 105% of the nominal speed of the gas generator of the turboshaft engine in a period of less than 10 seconds”, and paragraph [0024], “a mode, referred to as the normal reactivation mode, in which the turboshaft engine is rotated from the standby mode up to a speed in the range of between 80 and 105% of the nominal speed of the gas generator of the turboshaft engine in a period in the range of between 10 seconds and 60 seconds.”), and the rapid activation mode…allowing a time of activation of the second turboshaft engine of between 5 seconds and 15 seconds (See at least paragraph [0023], “a mode, referred to as the rapid reactivation mode, in which the turboshaft engine is rotated from the standby mode up to a speed in the range of between 80 and 105% of the nominal speed of the gas generator of the turboshaft engine in a period of less than 10 seconds.”).
Vallart does not explicitly disclose, however, Vhora, in the same field of endeavor, teaches and the use of a fuel feed-rate law… (See at least paragraph [0029], “The gas path model 15 is shown in FIG. 2 in detail, where external inputs 21, including customer bleeds, shaft power load, and starter torque are sent from sensors (in external inputs 21) to a CMEM model 23, which is a Component Map-based Engine Model and it is a nonlinear model. The incoming engine data is stored and processed. Ambient operating inputs 25 including, for example, temperature, pressure, Mach number and speed. Engine control unit inputs 27 are input into CMEM 23, including fuel commands and surge bleed valve control elements. Residuals are calculated by comparing the actual data inputs (on a continuous or steady state basis) with model predicted values that model CMEM 23 has determined for a set of data that represents the operating conditions of the engine being monitored. Model outputs 29 are computed for all the operating conditions, including shaft speeds, all temperatures, all pressures and all air flows”, paragraph [0031], “A number of engines were evaluated in the field with the present invention and relevant sensor readings have been compared. EGT, N2, LOT, HOT and fuel flow are compared and the percent error is calculated. Percent error is defined as the Model data value minus the sensor data value times 100 and divided by the sensor data value. [(model-sensor)*100]/sensor. Shown in FIG. 4 is a comparison with field engine data, showing the % N2 error between model and engine data at 41, % EGT error at 43, and % fuel flow error at 45. FIG. 5 shows the Lube System Model data with % HOT error at 51 and % LOT error at 53”, and paragraph [0033], “Table 1 below shows the typical fault signatures for unit change in fault parameters. They represent the relative change at steady state for given step change in fault parameter. For example, if fault parameter is increased from 0 to 1% in HP turbine fault model, the over all N2 speed will be reduced by 0.6% from its nominal speed, fuel flow will be increased by 0.6%, etc. The fault model works in both steady state as well as dynamic conditions, and the numbers in the table are for illustration purposes. Usually very small fault conditions are not detectable due to operating noise and variations. However fault levels that warrant attention (e.g., 2% degradation of turbine) are easy enough to detect using the method of this invention. 1TABLE I Typical Fault Signatures (1% change) HP Turbine Bleed Band Lube System N2 -0.6% N2 +0.3% HOT +0.6% Fuel Flow +0.6% Fuel Flow +0.7% LOT +0.5% EGT + 12.degree. F. EGT + 9.degree. F. P3 - 0.5% P3 - 0.3%.”), …comprising the use of a fuel flow-rate law… (See at least paragraph [0029], “The gas path model 15 is shown in FIG. 2 in detail, where external inputs 21, including customer bleeds, shaft power load, and starter torque are sent from sensors (in external inputs 21) to a CMEM model 23, which is a Component Map-based Engine Model and it is a nonlinear model. The incoming engine data is stored and processed. Ambient operating inputs 25 including, for example, temperature, pressure, Mach number and speed. Engine control unit inputs 27 are input into CMEM 23, including fuel commands and surge bleed valve control elements. Residuals are calculated by comparing the actual data inputs (on a continuous or steady state basis) with model predicted values that model CMEM 23 has determined for a set of data that represents the operating conditions of the engine being monitored. Model outputs 29 are computed for all the operating conditions, including shaft speeds, all temperatures, all pressures and all air flows”, paragraph [0031], “A number of engines were evaluated in the field with the present invention and relevant sensor readings have been compared. EGT, N2, LOT, HOT and fuel flow are compared and the percent error is calculated. Percent error is defined as the Model data value minus the sensor data value times 100 and divided by the sensor data value. [(model-sensor)*100]/sensor. Shown in FIG. 4 is a comparison with field engine data, showing the % N2 error between model and engine data at 41, % EGT error at 43, and % fuel flow error at 45. FIG. 5 shows the Lube System Model data with % HOT error at 51 and % LOT error at 53”, and paragraph [0033], “Table 1 below shows the typical fault signatures for unit change in fault parameters. They represent the relative change at steady state for given step change in fault parameter. For example, if fault parameter is increased from 0 to 1% in HP turbine fault model, the over all N2 speed will be reduced by 0.6% from its nominal speed, fuel flow will be increased by 0.6%, etc. The fault model works in both steady state as well as dynamic conditions, and the numbers in the table are for illustration purposes. Usually very small fault conditions are not detectable due to operating noise and variations. However fault levels that warrant attention (e.g., 2% degradation of turbine) are easy enough to detect using the method of this invention. 1TABLE I Typical Fault Signatures (1% change) HP Turbine Bleed Band Lube System N2 -0.6% N2 +0.3% HOT +0.6% Fuel Flow +0.6% Fuel Flow +0.7% LOT +0.5% EGT + 12.degree. F. EGT + 9.degree. F. P3 - 0.5% P3 - 0.3%.”), …comprising the use of a specific starting system configured to use an assistance torque and a fuel flow-rate law… (See at least paragraph [0029], “The gas path model 15 is shown in FIG. 2 in detail, where external inputs 21, including customer bleeds, shaft power load, and starter torque are sent from sensors (in external inputs 21) to a CMEM model 23, which is a Component Map-based Engine Model and it is a nonlinear model. The incoming engine data is stored and processed. Ambient operating inputs 25 including, for example, temperature, pressure, Mach number and speed. Engine control unit inputs 27 are input into CMEM 23, including fuel commands and surge bleed valve control elements. Residuals are calculated by comparing the actual data inputs (on a continuous or steady state basis) with model predicted values that model CMEM 23 has determined for a set of data that represents the operating conditions of the engine being monitored. Model outputs 29 are computed for all the operating conditions, including shaft speeds, all temperatures, all pressures and all air flows”, paragraph [0031], “A number of engines were evaluated in the field with the present invention and relevant sensor readings have been compared. EGT, N2, LOT, HOT and fuel flow are compared and the percent error is calculated. Percent error is defined as the Model data value minus the sensor data value times 100 and divided by the sensor data value. [(model-sensor)*100]/sensor. Shown in FIG. 4 is a comparison with field engine data, showing the % N2 error between model and engine data at 41, % EGT error at 43, and % fuel flow error at 45. FIG. 5 shows the Lube System Model data with % HOT error at 51 and % LOT error at 53”, and paragraph [0033], “Table 1 below shows the typical fault signatures for unit change in fault parameters. They represent the relative change at steady state for given step change in fault parameter. For example, if fault parameter is increased from 0 to 1% in HP turbine fault model, the over all N2 speed will be reduced by 0.6% from its nominal speed, fuel flow will be increased by 0.6%, etc. The fault model works in both steady state as well as dynamic conditions, and the numbers in the table are for illustration purposes. Usually very small fault conditions are not detectable due to operating noise and variations. However fault levels that warrant attention (e.g., 2% degradation of turbine) are easy enough to detect using the method of this invention. 1TABLE I Typical Fault Signatures (1% change) HP Turbine Bleed Band Lube System N2 -0.6% N2 +0.3% HOT +0.6% Fuel Flow +0.6% Fuel Flow +0.7% LOT +0.5% EGT + 12.degree. F. EGT + 9.degree. F. P3 - 0.5% P3 - 0.3%.”).
Vallart and Vhora do not explicitly disclose, however, MARCONI, in the same field of endeavor, teaches …the accelerated activation mode… (See at least paragraph [0009]-[0011], “The applicants have proposed in particular the following two standby modes: a standby mode referred to as normal super-idling, in which the combustion chamber is ignited and the shaft of the gas generator rotates at a speed of between 20 and 60% of the nominal speed, a standby mode referred to as assisted super-idling, in which the combustion chamber is ignited and the shaft of the gas generator rotates, with mechanical assistance, at a speed of between 20 and 60% of the nominal speed.”), the accelerated activation mode…allowing an activation time of between 10 seconds and 1 minute (See at least paragraph [0008], “In FR1151717 and FR1359766, the applicants proposed methods for optimising the specific consumption of the turboshaft engines of a helicopter by the possibility of putting at least one turboshaft engine into a stable flight mode, referred to as continuous flight mode, and at least one turboshaft engine into a particular standby mode that it can leave in an emergency or in a normal manner, according to need. A transition out of standby mode is referred to as ‘normal’ when a change in the flight situation requires the turboshaft engine in standby to be activated, for example when the helicopter is going to transition from a cruise flight situation to a landing phase. A normal transition out of standby mode of this kind occurs over a period of between 10 seconds and 1 minute. A transition out of standby mode is referred to as ‘emergency’ when there is a failure or a power deficit in the active engine, or when the flight conditions suddenly become difficult. An emergency transition out of standby mode of this kind occurs over a period of less than 10 seconds” and paragraph [0009]-[0011], “The applicants have proposed in particular the following two standby modes: a standby mode referred to as normal super-idling, in which the combustion chamber is ignited and the shaft of the gas generator rotates at a speed of between 20 and 60% of the nominal speed, a standby mode referred to as assisted super-idling, in which the combustion chamber is ignited and the shaft of the gas generator rotates, with mechanical assistance, at a speed of between 20 and 60% of the nominal speed.”).
Vallart, Vhora, and MARCONI do not explicitly disclose, however, BEDRINE, in the same field of endeavor, teaches the normal activation mode comprising a thermal stabilisation at a predefined temperature of the second turboshaft engine (See at least paragraph [0021], “In a second aspect of the method for quickly stopping the helicopter rotor, the rotation of the extinguished gas generator by said at least one electrical machine can be maintained as long as at least one measured temperature of the gas generator is greater than a first threshold and/or the absolute value of a measured temperature gradient is greater than a second threshold”, paragraph [0022], “In one aspect, the rotation of the extinguished gas generator can thus be maintained for a predefined ventilation period, but shut off before the expiration of this ventilation period as soon as the measured temperature of the gas generator is less than the first threshold”, paragraph [0049], “In a second step 110 of the method, an electronic control unit determines whether the thermal stabilization phase of the gas generator 12 of the turbomachine 4 has already taken place. If it has already taken place, in a step 115 the control unit transmits a signal to control the extinction of the gas generator 12, following which the pilot commands the stop of the main rotor 2 by the application of a brake on the shaft 6 of the main rotor 2”, and paragraph [0050], “If, on the other hand, no thermal stabilization phase is detected, the control unit commands the extinction of the combustion chamber of the gas generator 12 of the turbomachine 4 in a step 120, then the braking of the main rotor 2 in a step 130, and the ventilation of the gas generator 12 of the turbomachine in a step 140. The ventilation of the gas generator 12 is accomplished by the rotation of the gas generator 12 by an electrical machine which is powered by the electrical network of the helicopter. The electrical network of the helicopter can be coupled to one or more battery(ies), and possibly to at least one auxiliary power unit.” The thermal stabilization phase is controlled using a measured temperature of the gas generator relative to a first threshold, which corresponds to a predefined temperature criterion for the engine.), …allowing an activation time of between 1 and 3 minutes (See at least paragraph [0005], “The duration of this thermal stabilization phase is generally predefined depending on the type of engine and is typically comprised between 30 seconds and 2 minutes.”).
Thus, it would have been obvious to one of ordinary skill in the art before the effective filing date to combine the invention of Vallart with the teachings of Vhora, Marconi, and BEDRINE such that the helicopter propulsion system of Vallart is further configured to compare operating parameters of the turboshaft engine with equivalent parameters of a model representing a healthy turboshaft engine; to utilize a fuel feed-rate law; and utilize a specific starting system configured to use an assistance torque and a fuel flow-rate law, as taught by Vhora (See paragraph [0029], [0031]-[0033].), to utilize an accelerated activation mode having a second activation duration, the accelerated activation mode also being favoured over the rapid activation mode, and utilize the accelerated activation mode…allowing an activation time of between 10 seconds and 1 minute, as taught by Marconi (See paragraphs [0008]-[0013].), and the normal activation mode comprises a step of thermal stabilisation at a predefined temperature of the assistance engine and is favoured when a pilot of the aircraft gives the instruction for this to the assistance engine, and allowing an activation time of between 1 and 3 minutes, as taught by BEDRINE (See paragraph [0005], [0021], [0022], [0048]-[0050].), with a reasonable expectation of success. The motivation for doing so would be to detect existing engine performance-related faults using a nonlinear engine model in conjunction with operational engine data, as taught by Vhora (See paragraph [0016].), to remedy the operating-temperature disadvantage of the normal super-idling mode while maintaining the engine in a standby mode, as taught by Marconi (See paragraph [0012], [0013].), and to protect the engine against mechanical stresses associated with strong variations in power demand during thermal stabliziation, as taught by BEDRINE (See paragraph [0004].).
Conclusion
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/JEWEL A KUNTZ/Examiner, Art Unit 3666
/ANNE MARIE ANTONUCCI/Supervisory Patent Examiner, Art Unit 3666
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