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 .
Status of Claims
This Office Action is in response to the application filed on 22 October 2025. Claims 1-9, 11-17, and 19 are presently pending and are presented for examination. Claims 10, 18, and 20 are cancelled.
Response to Amendments
In response to Applicant’s amendments dated 22 October 2025, Examiner accepts the amendments to the specification; withdraws the previous objections to the drawings; withdraws the previous 35 U.S.C. 112(b) rejections; withdraws the previous 35 U.S.C. 101 rejections; and withdraws the previous prior art rejections.
Response to Arguments
Applicant's arguments, see Remarks, filed 22 October 2025, have been fully considered but they are not persuasive or moot.
Applicant argues, see Remarks, pg. 8-11, that the amended structural features of an aircraft and instruments within the aircraft make the claims subject matter eligible. However, the structural features and instruments are claimed generically and are operating in their ordinary capacity and do not use the judicial exception (mathematical concepts and mental processes) in a manner that imposes a meaningful limit on the judicial exception. Additionally, the dependent claims are directed towards describing additional aspects of the judicial exceptions and the last office action describes what those additional aspects of the limitations are in the dependent claims. For these reasons, examiner is unpersuaded and maintains the corresponding rejections. For more information on how the claims are not subject matter eligible, see the Claim Rejections - 35 USC § 101 section, below.
Applicant argues, see Remarks, pg. 11-13, that US-4968879-A (“Clark”) fails to disclose “determining an aircraft parameter indicative of a position of a longitudinal axis of the aircraft with respect to a horizon line during a flight of the aircraft…determining an offset angle between the longitudinal axis of the aircraft and the wind parameter…” recited in claim 1. However, Clark discloses determining wind flow characteristics and an aircraft angle of attack, which is an offset angle between a longitudinal axis of the aircraft with respect to the horizon (i.e., pitch) and the direction of wind flow (see Clark, col. 3, line 63 – col. 4, line 1). For these reasons, examiner is unpersuaded and maintains the corresponding rejections. For more information on how prior art teaches the limitations of claim 1, see the Claim Rejections - 35 USC § 102 section, below.
Applicant’s arguments, see Remarks, pg. 13-14, with respect to claim(s) 11 have been considered but are moot because the new ground of rejection does not rely on any reference applied in the prior rejection of record for any teaching or matter specifically challenged in the argument.
Applicant argues, see Remarks, pg. 15-17, that Clark fails to disclose “determining an offset angle between the longitudinal axis of the aircraft and the relative wind based on the aircraft parameter and the wind parameter” and “determining an offset angle…by subtracting the angle of the relative wind from the pitch attitude”. However, Clark discloses determining wind flow characteristics and an aircraft angle of attack, which is an offset angle between a longitudinal axis of the aircraft with respect to the horizon (i.e., pitch) and the direction of wind flow (see Clark, col. 3, line 63 – col. 4, line 1). Additionally, an angle of attack is the angle, or difference, between a reference line (i.e., longitudinal axis) on a body and the vector representing the relative motion between the body and the fluid in which it is moving (i.e., relative wind direction). For these reasons, examiner is unpersuaded and maintains the corresponding rejections.
The remaining arguments are essentially the same as those addressed above and/or below and are unpersuasive for at least the same reasons. Therefore, examiner is unpersuaded and maintains the corresponding rejections.
Claim Objections
Claim(s) 1, 11, and 16 is/are objected to because of the following informalities:
Claim 1: “determining an offset angle between the longitudinal axis of the aircraft and the wind parameter based on” should be “determining an offset angle between the longitudinal axis of the aircraft and the relative wind direction
Claim 11: “one or more processors and a memory storing one or instructions” should be “one or more processors and a memory storing
Claim 16: “determining an offset angle between the pitch attitude of the aircraft and the relative wind” should be “determining an offset angle between the pitch attitude of the aircraft and the angle of the relative wind”.
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.
Claim(s) 1-9, 11-17, and 19 is/are rejected under 35 U.S.C. 101 because the claimed invention is directed to an abstract idea without significantly more.
101 Analysis – Step 1
Claim(s) 1 and 16 is/are directed to a method (i.e., a process). Therefore, claims 1 and 16 are within at least one of the four statutory categories.
101 Analysis – Step 2A, Prong I
Regarding Prong I of the Step 2A analysis in the 2019 PEG, the claims are to be analyzed to determine whether they recite subject matter that falls within one of the following groups of abstract ideas: a) mathematical concepts, b) certain methods of organizing human activity, and/or c) mental processes.
Independent claim 1 includes limitations that recite an abstract idea (emphasized below) and will be used as a representative claim for the remainder of the 101 rejection. Claim 1 recites:
A method of operation, comprising:
providing an aircraft comprising a fuselage, an airfoil connected to the fuselage, a propulsion system, and a flight information system comprising one or more sensors and a controller;
determining an aircraft parameter indicative of a position of a longitudinal axis of the aircraft with respect to a horizon line during a flight of the aircraft;
determining a wind parameter indicative of a relative wind direction during the flight of the aircraft, the wind parameter determined based on a temperature distribution about a leading edge of a component of the aircraft, the temperature distribution comprising a plurality of measured temperatures measured by the one or more sensors; and
determining an offset angle between the longitudinal axis of the aircraft and the wind parameter based on the aircraft parameter and the wind parameter;
wherein the one or more sensors are disposed within the airfoil and outside of an external environment of the aircraft.
The examiner submits that the foregoing bolded limitation constitutes a “mathematical concept” because under its broadest reasonable interpretation, the claim covers mathematical calculations. Specifically, the “determining an offset angle…” limitation encompasses the mathematical calculation of an angle (calculation based on principles of geometry). Accordingly, the claim recites at least one abstract idea.
101 Analysis – Step 2A, Prong II
Regarding Prong II of the Step 2A analysis in the 2019 PEG, the claims are to be analyzed to determine whether the claim, as a whole, integrates the abstract into a practical application. As noted in the 2019 PEG, it must be determined whether any additional elements in the claim beyond the abstract idea integrate the exception into a practical application in a manner that imposes a meaningful limit on the judicial exception. The courts have indicated that additional elements merely using a computer to implement an abstract idea, adding insignificant extra solution activity, or generally linking use of a judicial exception to a particular technological environment or field of use do not integrate a judicial exception into a “practical application.”
In the present case, the additional limitations beyond the above-noted abstract idea are as follows (where the underlined portions are the “additional limitations” while the bolded portions continue to represent the “abstract idea”):
A method of operation, comprising:
providing an aircraft comprising a fuselage, an airfoil connected to the fuselage, a propulsion system, and a flight information system comprising one or more sensors and a controller;
determining an aircraft parameter indicative of a position of a longitudinal axis of the aircraft with respect to a horizon line during a flight of the aircraft;
determining a wind parameter indicative of a relative wind direction during the flight of the aircraft, the wind parameter determined based on a temperature distribution about a leading edge of a component of the aircraft,
the temperature distribution comprising a plurality of measured temperatures measured by the one or more sensors; and
determining an offset angle between the longitudinal axis of the aircraft and the wind parameter based on the aircraft parameter and the wind parameter;
wherein the one or more sensors are disposed within the airfoil and outside of an external environment of the aircraft.
For the following reason(s), the examiner submits that the above identified additional limitations do not integrate the above-noted abstract idea into a practical application. Regarding the additional limitations of “determining an aircraft parameter…” and “determining a wind parameter…”, the examiner submits that these limitations are insignificant extra-solution activities. In particular, the parameter determination limitations are recited at a high level of generality (i.e., as a general means of gathering aircraft and wind data for use in the “determining an offset angle…” step), and amounts to mere data gathering, which is a form of insignificant extra-solution activity. Regarding the additional limitations of “providing an aircraft comprising…”, “the temperature distribution comprising…”, and “wherein the one or more sensors are disposed…”, the examiner submits that these limitations are an attempt to generally link additional elements to a technological environment. In particular, the aircraft components, including the “one or more sensors” and the “controller” are recited at a high level of generality. The aircraft components, including the “one or more sensors” and the “controller”, are claimed generically and are operating in their ordinary capacity and do not use the judicial exception in a manner that imposes a meaningful limit on the judicial exception, such that the claim is more than a drafting effort designed to monopolize the exception.
Thus, taken alone, the additional elements do not integrate the abstract idea into a practical application. Further, looking at the additional limitation(s) as an ordered combination or as a whole, the limitation(s) add nothing that is not already present when looking at the elements taken individually. For instance, there is no indication that the additional elements, when considered as a whole, reflect an improvement in the functioning of a computer or an improvement to another technology or technical field, apply or use the above-noted judicial exception to effect a particular treatment or prophylaxis for a disease or medical condition, implement/use the above-noted judicial exception with a particular machine or manufacture that is integral to the claim, effect a transformation or reduction of a particular article to a different state or thing, or apply or use 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 not more than a drafting effort designed to monopolize the exception (MPEP § 2106.05). Accordingly, the additional limitation(s) do/does not integrate the abstract idea into a practical application because it does not impose any meaningful limits on practicing the abstract idea.
101 Analysis – Step 2B
Regarding Step 2B of the 2019 PEG, representative 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 to those discussed above with respect to determining that the claim does not integrate the abstract idea into a practical application. As discussed above with respect to integration of the abstract idea into a practical application, the additional limitations of “determining an aircraft parameter…” and “determining a wind parameter…”, the examiner submits that these limitations are insignificant extra-solution activities and the limitations of “providing an aircraft comprising…”, “the temperature distribution comprising…”, and “wherein the one or more sensors are disposed…”, the examiner submits that these limitations are an attempt to generally link additional elements to a technological environment.
Further, a conclusion that an additional element is insignificant extra-solution activity in Step 2A should be re-evaluated in Step 2B to determine if they are more than what is well-understood, routine, conventional activity in the field. The additional limitations of “determining an aircraft parameter…” and “determining a wind parameter…” are well-understood, routine, and conventional activities because the sensors are all conventional sensors mounted on the aircraft: aircraft parameter sensor(s) (para. 0043: “gyroscope 56 or any other suitable aircraft instrument”; see also para. 0008 and 0017) and wind parameter sensor(s) (para. 0021: “using one or more temperature sensors”; see also para. 0007, 0019-20, 0037-0040, and 0045). MPEP 2106.05(d)(II), and the cases cited therein, including Intellectual Ventures I, LLC v. Symantec Corp., 838 F.3d 1307, 1321 (Fed. Cir. 2016), TLI Communications LLC v. AV Auto. LLC, 823 F.3d 607, 610 (Fed. Cir. 2016), and OIP Techs., Inc., v. Amazon.com, Inc., 788 F.3d 1359, 1363 (Fed. Cir. 2015), indicate that mere collection or receipt of data over a network is a well‐understood, routine, and conventional function when it is claimed in a merely generic manner. Hence, independent claim 1 and analogous independent claim 16 are not patent subject matter eligible.
Dependent claim(s) 2-9, 17, and 19 do not recite any further limitations that cause the claim(s) to be patent subject matter eligible. Rather, the limitations of dependent claims are directed toward additional aspects of the judicial exception and/or well-understood, routine and conventional additional elements that do not integrate the judicial exception into a practical application. Specifically, they are directed towards additional aspects of the abstract idea, the additional elements linking the abstract idea to a technological environment, and the data gathering limitations: the types of angles determined (see claims 2, 4, and 19), the types of aircraft and wind data to be gathered (see claims 3, 6, and 17), and the sensors (type and implemented locations) used to gather the data (see claims 5, 7-9, and 19). Therefore, dependent claims 2-9, 17, and 19 are not patent subject matter eligible under the same rationale as provided for in the rejection of independent claims 1 and 16.
Therefore, claims 1-9, 16-17, and 19 are ineligible under 35 U.S.C. 101.
101 Analysis – Step 1
Claim(s) 11 is/are directed to an aircraft (i.e., a machine). Therefore, claim 11 is within at least one of the four statutory categories.
101 Analysis – Step 2A, Prong I
Regarding Prong I of the Step 2A analysis in the 2019 PEG, the claims are to be analyzed to determine whether they recite subject matter that falls within one of the following groups of abstract ideas: a) mathematical concepts, b) certain methods of organizing human activity, and/or c) mental processes.
Independent claim 11 includes limitations that recite an abstract idea (emphasized below). Claim 11 recites:
An aircraft, comprising:
a nose;
one or more sensors disposed within the nose of the aircraft and configured to measure a temperature at a plurality of measurement locations about a leading edge of the nose to determine a temperature distribution; and
one or more processors and a memory storing one or instructions, the one or more processors configured to execute the instructions stored on the memory to cause the one or more processors to:
determine a temperature distribution based on the temperature at the plurality of measurement locations measured via the one or more sensors;
identifying a high temperature location out of the plurality of measurement locations with a highest measured temperature in the temperature distribution; and
determining a parameter for the aircraft based on a position of the high temperature location about the leading edge of the nose of the aircraft;
wherein the nose comprises an exterior skin exposed to an external environment and a skin interior surface forming an internal cavity; and
wherein the one or more sensors are thermally coupled to the skin interior surface within the internal cavity of the nose of the aircraft.
The examiner submits that the foregoing bolded limitations constitute a “mental process” because under its broadest reasonable interpretation, the claim covers performance of the limitation in the human mind. Specifically, the “determine a temperature distribution…”, the “identifying a high temperature location…”, and the “determining a parameter for the aircraft…” limitations encompass a person observing temperature data of an aircraft component and making an observation, evaluation, or judgement about the temperature data and then using those observations, evaluations, or judgements to determine an aircraft parameter. Accordingly, the claim recites at least one abstract idea.
101 Analysis – Step 2A, Prong II
Regarding Prong II of the Step 2A analysis in the 2019 PEG, the claims are to be analyzed to determine whether the claim, as a whole, integrates the abstract into a practical application. As noted in the 2019 PEG, it must be determined whether any additional elements in the claim beyond the abstract idea integrate the exception into a practical application in a manner that imposes a meaningful limit on the judicial exception. The courts have indicated that additional elements merely using a computer to implement an abstract idea, adding insignificant extra solution activity, or generally linking use of a judicial exception to a particular technological environment or field of use do not integrate a judicial exception into a “practical application.”
In the present case, the additional limitations beyond the above-noted abstract idea are as follows (where the underlined portions are the “additional limitations” while the bolded portions continue to represent the “abstract idea”):
An aircraft, comprising:
a nose;
one or more sensors disposed within the nose of the aircraft and configured to measure a temperature at a plurality of measurement locations about a leading edge of the nose to determine a temperature distribution; and
one or more processors and a memory storing one or instructions, the one or more processors configured to execute the instructions stored on the memory to cause the one or more processors to:
determine a temperature distribution based on the temperature at the plurality of measurement locations measured via the one or more sensors;
identifying a high temperature location out of the plurality of measurement locations with a highest measured temperature in the temperature distribution; and
determining a parameter for the aircraft based on a position of the high temperature location about the leading edge of the nose of the aircraft;
wherein the nose comprises an exterior skin exposed to an external environment and a skin interior surface forming an internal cavity; and
wherein the one or more sensors are thermally coupled to the skin interior surface within the internal cavity of the nose of the aircraft.
For the following reason(s), the examiner submits that the above identified additional limitations do not integrate the above-noted abstract idea into a practical application. Regarding the additional limitations of the aircraft “nose…”, “one or more sensors…”, “wherein the nose comprises…” and “wherein the one or more sensors…”, the examiner submits that these limitations are an attempt to generally link additional elements to a technological environment. In particular, the aircraft components, including the “nose” and the “one or more sensors” are recited at a high level of generality. The aircraft components, including the “nose” and the “one or more sensors”, are claimed generically and are operating in their ordinary capacity and do not use the judicial exception in a manner that imposes a meaningful limit on the judicial exception, such that the claim is more than a drafting effort designed to monopolize the exception. Regarding the additional limitations of “determining” and “identifying” by the “one or more processors and a memory” are recited at a high level of generality and merely automates the “determining” and “identifying” limitations, therefore acting as a generic computer to perform the abstract idea. The “one or more processors and a memory” is claimed generically and is operating in its ordinary capacity and does not use the judicial exception in a manner that imposes a meaningful limit on the judicial exception, such that the claim is more than a drafting effort designed to monopolize the exception. The additional limitations are no more than mere instructions to apply the exception using a computer (the “one or more processors and a memory”).
Thus, taken alone, the additional elements do not integrate the abstract idea into a practical application. Further, looking at the additional limitation(s) as an ordered combination or as a whole, the limitation(s) add nothing that is not already present when looking at the elements taken individually. For instance, there is no indication that the additional elements, when considered as a whole, reflect an improvement in the functioning of a computer or an improvement to another technology or technical field, apply or use the above-noted judicial exception to effect a particular treatment or prophylaxis for a disease or medical condition, implement/use the above-noted judicial exception with a particular machine or manufacture that is integral to the claim, effect a transformation or reduction of a particular article to a different state or thing, or apply or use 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 not more than a drafting effort designed to monopolize the exception (MPEP § 2106.05). Accordingly, the additional limitation(s) do/does not integrate the abstract idea into a practical application because it does not impose any meaningful limits on practicing the abstract idea.
101 Analysis – Step 2B
Regarding Step 2B of the 2019 PEG, representative independent claim 11 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 to those discussed above with respect to determining that the claim does not integrate the abstract idea into a practical application. As discussed above with respect to integration of the abstract idea into a practical application, the additional limitations of a “nose”, “one or more sensors”, “wherein the nose comprises…”, and “wherein the one or more sensors…” the examiner submits that these limitations are an attempt to generally link additional elements to a technological environment. The limitations of “one or more processors and a memory”, the examiner submits that these limitations are no more than mere instructions to apply the exception using a computer. Hence, independent claim 11 is not patent subject matter eligible.
Dependent claim(s) 12-15 do not recite any further limitations that cause the claim(s) to be patent subject matter eligible. Rather, the limitations of dependent claims are directed toward additional aspects of the judicial exception and/or well-understood, routine and conventional additional elements that do not integrate the judicial exception into a practical application. Specifically, they are directed towards additional aspects of the “determining a parameter” limitation: the type of parameter determined (see claim 12-13) and the type of additional determined data based on the determined parameter (see claim 14-15). Therefore, dependent claims 12-15 are not patent subject matter eligible under the same rationale as provided for in the rejection of independent claim 11.
Therefore, claims 11-15 are ineligible under 35 U.S.C. 101.
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.
Claim(s) 1-4 is/are rejected under 35 U.S.C. 102(a)(1) as being anticipated by US-4968879-A, hereinafter “Clark” (previously of record).
Regarding claim 1, Clark discloses A method of operation (Clark, col. 1, lines 13-14: “This invention relates to a device and method for measuring aircraft flight parameters...”), comprising:
providing an aircraft comprising a fuselage, an airfoil connected to the fuselage, a propulsion system, and a flight information system comprising one or more sensors and a controller (Clark, FIG. 8-9: [i.e., providing an aircraft comprising a fuselage, an airfoil connected to the fuselage, a propulsion system]; col. 3, lines 22-27: “Radiation sensors 26 and 28 [i.e., a flight information system comprising one or more sensors] transmit the electrical data to the aircraft control center, such as computer 30 [i.e., a controller]. The data are analyzed in the computer 30 to determine aircraft flight parameters such as velocity and attitude. The aircraft flight parameters are used by either the computer, the pilot, or both, to control the aircraft.”);
determining an aircraft parameter indicative of a position of a longitudinal axis of the aircraft with respect to a horizon line during a flight of the aircraft (Clark, col. 4, lines 37-49: “The yaw, roll and other attitudes may be calculated [i.e., determining an aircraft parameter] using appropriately placed thermal radiators and sensors. For each attitude to be sensed, two thermal radiators are placed on either side of the axis of interest. For each aircraft member or aircraft flight parameter of interest, the radiation emitted by two or more sensors appropriately placed can be compared and analyzed to determine the flight status, as has been described with respect to determining the pitch of the nose [i.e., indicative of a position of a longitudinal axis of the aircraft with respect to a horizon line during a flight of the aircraft]. For example, to measure the pitch of a member, thermal radiators may be positioned on an upper and lower surface, and to measure yaw of a member, thermal radiators may be positioned on left and right surfaces, etc.”);
determining a wind parameter indicative of a relative wind direction during the flight of the aircraft, the wind parameter determined based on a temperature distribution about a leading edge of a component of the aircraft, the temperature distribution comprising a plurality of measured temperatures measured by the one or more sensors (Clark, col. 3, lines 7-10: ” The quantity of thermal radiation emitted to the sensors 26 and 28 [i.e., the temperature distribution comprising a plurality of measured temperatures measured by the one or more sensors] is detected, and an electrical signal is generated corresponding to the quantity of thermal radiation received.”; col. 3, lines 47-55: “The relationship between the angle of attack of the skin surface and the heat flux of the skin surface [i.e., based on a temperature distribution] is illustrated in FIG. 5. FIG. 5 is based on a leading edge of an aircraft [i.e., leading edge of a component of the aircraft] having a radius of four inches, a constant air density, and a constant aircraft speed as the angle of the skin surface varies. When the angle of the skin surface with respect to the airflow is between 5° and approximately 20° [i.e., relative wind direction during the flight of the aircraft], the heat flux of the skin surface due to air friction is generally linear, as can be seen from FIG. 5.”; col. 4, lines 50-56: “If the aircraft member 10 does not extend uniformly with respect to a desired plane of measurement, such as in a wing member, the nose of some aircraft, etc., the attitude analysis must take into account differences in angles of the surfaces with respect to each other, differences in air density, and airflow patterns [i.e., determining a wind parameter] over the different surfaces, and possibly numerous other factors.”); and
determining an offset angle between the longitudinal axis of the aircraft and the wind parameter based on the aircraft parameter and the wind parameter (Clark, col. 3, line 37-46: “As the aircraft travels through the air, the friction from the air causes significant heating of the exposed skin surfaces of the aircraft, particularly leading edges. The temperature of the skin of an aircraft flying at high supersonic or hypersonic speeds is proportional to the angle of the skin surface with respect to the flow stream of air [i.e., relative wind direction; wind parameter] for a given velocity of the aircraft at a given air density. Further, at a given air density and angle of attack [i.e., an offset angle between the longitudinal axis of the aircraft and the wind parameter based on the aircraft parameter and the wind parameter], the heating due to air friction is proportional to the velocity of aircraft in the air.”; col. 3, line 63-col. 4 line 1: “Therefore, any differences in the thermal radiation between the members correspond to differences in angle of the skin with respect to the airflow. For example, assuming θ is 10° and ɑ [i.e., aircraft parameter] is zero, then the angle of attack [i.e., determining an offset angle] of the nose of the plane is zero and the radiation emitted by the respective members 20 and 22 will be equal.”);
wherein the one or more sensors are disposed within the airfoil and outside of an external environment of the aircraft (Clark, FIGs. 1 and 4; col. 2, lines 56-59: “While the member 10 as shown is the nose of the aircraft, any other members, such as a curved wing surface [i.e., within the airfoil], rudder, elevator or other aircraft skin surface member, may be used.”; col. 3, lines 4-7: “Fiberoptic cables 24 are mounted directly adjacent the thermal radiation members 20 and 22 for guiding the thermal radiation to respective radiation detectors 26 and 28 [i.e., wherein the one or more sensors are disposed within the airfoil and outside of an external environment of the aircraft].”; col. 5, line 63 - col. 6, line 1: “Alternatively, as shown in FIG. 4, the thermal radiation member 20 may be mounted directly behind skin surface 12. Heat is transferred by skin surface 12 to thermal radiation member 20. The thermal radiation member 20 of the embodiment of FIG. 4 is mounted in direct contact with the skin member 12, but not in contact with the external atmosphere [i.e., wherein the one or more sensors are disposed within the airfoil and outside of an external environment of the aircraft].”).
Regarding claim 2, Clark discloses The method of claim 1,
wherein the offset angle comprises an angle of attack of the aircraft (Clark, col. 3, lines 58-61: “The angle of attack can be determined by measuring the difference between the thermal radiation emitted by members 20 and 22.”).
Regarding claim 3, Clark discloses The method of claim 2,
wherein the aircraft parameter comprises a pitch attitude of the aircraft between the longitudinal axis of the aircraft and the horizon line (Clark, col. 4, lines 37-49: “The yaw, roll and other attitudes may be calculated using appropriately placed thermal radiators and sensors. For each attitude to be sensed, two thermal radiators are placed on either side of the axis of interest. For each aircraft member or aircraft flight parameter of interest, the radiation emitted by two or more sensors appropriately placed can be compared and analyzed to determine the flight status, as has been described with respect to determining the pitch of the nose [i.e., aircraft parameter comprises a pitch attitude of the aircraft and the horizon line]. For example, to measure the pitch of a member, thermal radiators may be positioned on an upper and lower surface, and to measure yaw of a member, thermal radiators may be positioned on left and right surfaces, etc.”); and
the wind parameter comprises a flight path angle between the relative wind direction and the horizon line (Clark, col. 1, line 66 - col. 2, line 2: “It is therefore an object of this invention to provide an apparatus for rapidly measuring various flight parameters [i.e., a flight path angle] of an aircraft traveling at high supersonic or hypersonic speeds, including attitude, velocity, angle of attack and the like.”; Note: To one of ordinary skill in the art, at the time of the application, the following flight parameters, angle of attack, pitch angle, and flight path angle, are obvious variants and their mathematical relationship is well-known (pitch angle = angle of attack + flight path angle).).
Regarding claim 4, Clark discloses The method of claim 1,
wherein the offset angle comprises a sideslip angle of the aircraft (Clark, col. 2, line 62-68: “When member 10 is the nose of the aircraft, the angle of attack, α, illustrated in FIG. 1, is the pitch of the aircraft as the aircraft climbs or dives. Other aircraft attitudes, such as the yaw [i.e., the offset angle comprises a sideslip angle of the aircraft] and roll, may also be measured according to this invention following the same principles used in measuring the pitch.”).
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) 5 is/are rejected under 35 U.S.C. 103 as being unpatentable over Clark, as applied to claim 1 above, and further in view of “A low cost MEMS and complementary filter based attitude heading reference system (AHRS) for low speed aircraft”, hereinafter “Islam” (previously of record).
Regarding claim 5, Clark discloses The method of claim 1, but does not appear to explicitly disclose the following:
wherein the aircraft parameter is determined using a gyroscope onboard the aircraft.
However, in the same field of endeavor, Islam teaches:
wherein the aircraft parameter is determined using a gyroscope onboard the aircraft (Islam, Abstract: “AHRS is an important system that provides the 3-dimensional orientation (roll, pitch and yaw or heading) of an aircraft from which the flying performance can be evaluated.”; col. 1, last paragraph: “An IMU generally comprises of three axis gyroscopes, accelerometers and magnetometers for measuring the 3D orientation (roll, pitch and yaw angle).”; col. 2, paragraph 3: “This paper proposes a standalone AHRS utilizing low cost MEMS based three axis accelerometers, gyroscopes and magnetometers using a complementary filter to determine the flying performance of a low speed aircraft.”).
Therefore, it would have been obvious to one of ordinary skill in the art, before the effective filing date of the claimed invention and with a reasonable likelihood of success to modify the invention disclosed by Clark, with the concept of using at least one gyroscope to determine orientation of an aircraft, taught by Islam, in order to maintain control of the aircraft based on accurate estimates of speed and orientation (Islam, Abstract: “AHRS is an important system that provides the 3-dimensional orientation (roll, pitch and yaw or heading) of an aircraft from which the flying performance can be evaluated.”).
Claim(s) 6-9, 11-12, and 14-15 is/are rejected under 35 U.S.C. 103 as being unpatentable over Clark, as applied to claim 1 above, and further in view of CN-106679932-A, hereinafter “Li” (previously of record).
Regarding claim 6, Clark discloses The method of claim 1,
wherein the determining of the wind parameter comprises measuring a temperature at each of a plurality of measurement locations about the leading edge of the component of the aircraft to determine the temperature distribution (Clark, col. 3, lines 7-10: ” The quantity of thermal radiation emitted to the sensors 26 and 28 is detected, and an electrical signal is generated corresponding to the quantity of thermal radiation received [i.e., comprises measuring a temperature at each of a plurality of measurement locations about the leading edge of the component of the aircraft to determine the temperature distribution].”; col. 3, lines 47-55: “The relationship between the angle of attack of the skin surface and the heat flux of the skin surface is illustrated in FIG. 5. FIG. 5 is based on a leading edge of an aircraft having a radius of four inches, a constant air density, and a constant aircraft speed as the angle of the skin surface varies. When the angle of the skin surface with respect to the airflow is between 5° and approximately 20°, the heat flux of the skin surface due to air friction is generally linear, as can be seen from FIG. 5.”; col. 4, lines 50-56: “If the aircraft member 10 does not extend uniformly with respect to a desired plane of measurement, such as in a wing member, the nose of some aircraft, etc., the attitude analysis [i.e., determining a wind parameter] must take into account differences in angles of the surfaces with respect to each other, differences in air density, and airflow patterns over the different surfaces, and possibly numerous other factors.”);
Clark does not appear to explicitly disclose the following:
identifying a high temperature location out of the plurality of measurement locations with a highest measured temperature in the temperature distribution; and determining the wind parameter based on a position of the high temperature location about the leading edge of the component of the aircraft.
However, in the same field of endeavor, Li teaches:
identifying a high temperature location out of the plurality of measurement locations with a highest measured temperature in the temperature distribution (translated document of Li, para. 30: “Low speed - sub-acoustic airfoil angle of attack, including low-speed - sub-acoustic airfoil, airfoil in the air, low speed - sub-acoustic airfoil sensor array layout, low speed - sub-acoustic airfoil leading edge temperature distribution, And the relationship between the angle of attack, in the course of the movement of the aircraft, due to the direction of the flow direction of the thermal element measured film surface temperature Tc will be the largest [i.e., a highest measured temperature in the temperature distribution], the remaining position with the flow of air, Tc will gradually reduce the formation of different angles”); and
determining the wind parameter based on a position of the high temperature location about the leading edge of the component of the aircraft (translated document of Li, para. 30: “Low speed - sub-acoustic airfoil angle of attack, including low-speed - sub-acoustic airfoil, airfoil in the air, low speed - sub-acoustic airfoil sensor array layout, low speed - sub-acoustic airfoil leading edge [i.e., about the leading edge of the component of the aircraft] temperature distribution, And the relationship between the angle of attack [i.e., determining the wind parameter], in the course of the movement of the aircraft, due to the direction of the flow direction of the thermal element measured film surface temperature Tc will be the largest [i.e., based on a position of the high temperature location], the remaining position with the flow of air, Tc will gradually reduce the formation of different angles”).
Therefore, it would have been obvious to one of ordinary skill in the art, before the effective filing date of the claimed invention and with a reasonable likelihood of success to modify the invention disclosed by Clark, with the concept of measuring temperature at a plurality of locations about a leading edge of an aircraft component, identifying the highest temperature location and determining a wind parameter based on the location of the highest temperature location, taught by Li, in order to accurately determine an attitude (like the angle of attack) of an aircraft (translated document of Li, para. 50: “…the surface temperature Tc of the thin film layer of the thermal sensor measured at the direction facing the incoming flow will be the largest, while the rest of the positions will gradually decrease with the flow of the airflow, forming different angles of attack. Corresponding to different low speed-subsonic airfoil leading edge temperature distribution, with the change of the angle of attack, different sensitive units of the micro flexible hot film sensor array will measure the maximum value, and the peak value will also move with the change of the angle of attack, as shown in The relationship between the extreme temperature position and the angle of attack; when the angle of attack is 0°, the extreme temperature is at the origin; when the angle of attack is positive, the extreme temperature is at a position on the negative semi-axis; when the angle of attack is negative, the temperature The extreme value is located at a certain position on the positive semi-axis, so that the change in the angle of attack can be converted into the position change of the temperature extreme value along the latitude direction, and the data collected in step 2) can be used to fit the temperature extreme value of the low-subsonic airfoil The relationship curve between the position and the angle of attack can be used to determine the angle of attack by the position of the temperature extreme value measured in real time during the flight…”).
Regarding claim 7, Clark and Li teach The method of claim 6, and Clark further discloses the following:
wherein the component of the aircraft comprises a skin (Clark, col. 2, lines 15-20: “. A fiberoptic cable is coupled between the thermal radiation sensor and the skin of the aircraft for transferring the emitted radiation to the sensor. The radiation emitted by the skin at different locations of the aircraft is analyzed to determine various aircraft attitudes.”) and
a cavity, an interior surface of the skin forms a peripheral boundary of the cavity within the component of the aircraft (Clark, FIGs. 2 and 4; col. 5, lines 15-18: “As illustrated in FIG. 2, the fiberoptic cable 24 may be mounted in a cavity formed in the skin member 12, the skin member itself being thermal radiation member 20.”), and
an exterior surface of the skin is exposed to an environment external to the aircraft (Clark, FIGs. 2-4; col. 1, lines 32-35: “One difficulty in measuring flight parameters of supersonic and hypersonic vehicles is the severe thermal environment [i.e., exposed to an environment external to the aircraft] and broad range of temperature fluctuations at the skin of the aircraft [i.e., an exterior surface of the skin].”); and
the temperature is measured at each of the plurality of measurement locations using one or more temperature sensors connected to the skin at the interior surface (Clark, FIGs. 2-4; col. 2, lines 13-15: “These and other objects of the invention are accomplished by providing a thermal radiation sensor [i.e., temperature is measured…using one or more temperature sensors] for sensing thermal radiation from the skin of the aircraft [i.e., connected to the skin at the interior surface]. A fiberoptic cable is coupled between the thermal radiation sensor and the skin of the aircraft for transferring the emitted radiation to the sensor. The radiation emitted by the skin at different locations [i.e., plurality of measurement locations] of the aircraft is analyzed to determine various aircraft attitudes.”).
Regarding claim 8, Clark and Li teach The method of claim 6, and Clark further discloses the following:
wherein the component of the aircraft comprises a skin (Clark, col. 2, lines 15-20: “. A fiberoptic cable is coupled between the thermal radiation sensor and the skin of the aircraft for transferring the emitted radiation to the sensor. The radiation emitted by the skin at different locations of the aircraft is analyzed to determine various aircraft attitudes.”) and
a cavity, an interior surface of the skin forms a peripheral boundary of the cavity within the component of the aircraft (Clark, FIGs. 2 and 4; col. 5, lines 15-18: “As illustrated in FIG. 2, the fiberoptic cable 24 may be mounted in a cavity formed in the skin member 12, the skin member itself being thermal radiation member 20.”), and
an exterior surface of the skin is exposed to an environment external to the aircraft (Clark, FIGs. 2-4; col. 1, lines 32-35: “One difficulty in measuring flight parameters of supersonic and hypersonic vehicles is the severe thermal environment [i.e., exposed to an environment external to the aircraft] and broad range of temperature fluctuations at the skin of the aircraft [i.e., an exterior surface of the skin].”); and
the temperature is measured at each of the plurality of measurement locations using one or more temperature sensors disposed within the cavity and thermally coupled to the skin (Clark, FIGs. 2-4; col. 2, lines 13-15: “These and other objects of the invention are accomplished by providing a thermal radiation sensor [i.e., temperature is measured…using one or more temperature sensors] for sensing thermal radiation from the skin of the aircraft. A fiberoptic cable is coupled between the thermal radiation sensor and the skin of the aircraft for transferring the emitted radiation to the sensor. The radiation emitted by the skin at different locations [i.e., plurality of measurement locations] of the aircraft is analyzed to determine various aircraft attitudes.”; col. 5, line 63 – col. 6, line 1: “Alternatively, as shown in FIG. 4, the thermal radiation member 20 may be mounted directly behind skin surface 12. Heat is transferred by skin surface 12 to thermal radiation member 20. The thermal radiation member 20 of the embodiment of FIG. 4 is mounted in direct contact with the skin member 12 [i.e., disposed within the cavity and thermally coupled to the skin], but not in contact with the external atmosphere. ”).
Regarding claim 9, Clark and Li teach The method of claim 6, and Clark further discloses the following:
wherein the temperature is measured at each of the plurality of measurement locations using one or more temperature sensors disposed within the component of the aircraft (Clark, FIGs. 2-4; col. 2, lines 15-20: “A fiberoptic cable is coupled between the thermal radiation sensor [i.e., temperature is measured…using one or more temperature sensors] and the skin of the aircraft [i.e., disposed within the component of the aircraft] for transferring the emitted radiation to the sensor. The radiation emitted by the skin at different locations [i.e., plurality of measurement locations] of the aircraft is analyzed to determine various aircraft attitudes.”).
Regarding claim 11, Clark discloses An aircraft (Clark, FIG. 8-9), comprising:
a nose (Clark, FIG. 1 [i.e., a nose]; col. 2, lines 55-59: “An aircraft member, designated generally as 10, includes exposed skin surfaces 12 and 14. While the member 10 as shown is the nose of the aircraft [i.e., a nose], any other members, such as a curved wing surface [i.e., within the airfoil], rudder, elevator or other aircraft skin surface member, may be used.”;);
one or more sensors disposed within the nose of the aircraft (Clark, FIG. 1-4; col. 3, lines 4-7: “Fiberoptic cables 24 are mounted directly adjacent the thermal radiation members 20 and 22 for guiding the thermal radiation to respective radiation detectors 26 and 28 [i.e., one or more sensors disposed within the nose of the aircraft].”; col. 5, line 63 - col. 6, line 1: “Alternatively, as shown in FIG. 4, the thermal radiation member 20 may be mounted directly behind skin surface 12. Heat is transferred by skin surface 12 to thermal radiation member 20. The thermal radiation member 20 of the embodiment of FIG. 4 is mounted in direct contact with the skin member 12, but not in contact with the external atmosphere [i.e., one or more sensors disposed within the nose of the aircraft].”) and
configured to measure a temperature at a plurality of measurement locations about a leading edge of the nose to determine a temperature distribution (Clark, FIG. 1-4; col. 3, lines 7-10: ” The quantity of thermal radiation emitted to the sensors 26 and 28 [i.e., configured to measure a temperature at a plurality of measurement locations] is detected, and an electrical signal is generated corresponding to the quantity of thermal radiation received.”; col. 3, lines 47-55: “The relationship between the angle of attack of the skin surface and the heat flux of the skin surface is illustrated in FIG. 5. FIG. 5 is based on a leading edge of an aircraft [i.e., about a leading edge of the nose to determine a temperature distribution] having a radius of four inches, a constant air density, and a constant aircraft speed as the angle of the skin surface varies. When the angle of the skin surface with respect to the airflow is between 5° and approximately 20° [i.e., relative wind direction during the flight of the aircraft], the heat flux of the skin surface due to air friction is generally linear, as can be seen from FIG. 5.”); and
one or more processors and a memory storing one or instructions, the one or more processors configured to execute the instructions stored on the memory to cause the one or more processors to (Clark, col. 3, lines 21-26: “Radiation sensors 26 and 28 transmit the electrical data to the aircraft control center, such as computer 30 [i.e., one or more processors and a memory storing one or instructions]. The data are analyzed in the computer 30 [i.e., the one or more processors configured to execute the instructions stored on the memory] to determine aircraft flight parameters such as velocity and attitude. The aircraft flight parameters are used by either the computer, the pilot, or both, to control the aircraft.”):
determine a temperature distribution based on the temperature at the plurality of measurement locations measured via the one or more sensors (Clark, FIG. 1-4; col. 3, lines 7-10: ” The quantity of thermal radiation emitted to the sensors 26 and 28 [i.e., a temperature distribution based on the temperature at the plurality of measurement locations measured via the one or more sensors] is detected, and an electrical signal is generated corresponding to the quantity of thermal radiation received.”)…
…wherein the nose comprises an exterior skin exposed to an external environment and a skin interior surface forming an internal cavity (Clark, FIG. 1-4: fiber optic and sensor 26 mounted in cavity; col. 2, lines 55-56: “An aircraft member, designated generally as 10, includes exposed skin surfaces 12 and 14 [i.e., nose comprises an exterior skin exposed to an external environment].”; col. 5, lines 14-18: “The fiberoptic cable may be mounted a variety of ways with respect to the thermal radiator. As illustrated in FIG. 2, the fiberoptic cable 24 may be mounted in a cavity formed in the skin member 12 [i.e., skin interior surface forming an internal cavity], the skin member itself being thermal radiation member 20.”; col. 5, line 63 – col. 6, line 1: “Alternatively, as shown in FIG. 4, the thermal radiation member 20 may be mounted directly behind skin surface 12. Heat is transferred by skin surface 12 to thermal radiation member 20. The thermal radiation member 20 of the embodiment of FIG. 4 is mounted in direct contact with the skin member 12, but not in contact with the external atmosphere.”); and
wherein the one or more sensors are thermally coupled to the skin interior surface within the internal cavity of the nose of the aircraft (Clark, FIG. 1-4: fiber optic and sensor 26 mounted in cavity; col. 5, line 63 – col. 6, line 1: “Alternatively, as shown in FIG. 4, the thermal radiation member 20 may be mounted directly behind skin surface 12. Heat is transferred by skin surface 12 to thermal radiation member 20. The thermal radiation member 20 of the embodiment of FIG. 4 is mounted in direct contact with the skin member 12, but not in contact with the external atmosphere.”).
Clark does not appear to explicitly disclose the following:
identifying a high temperature location out of the plurality of measurement locations with a highest measured temperature in the temperature distribution; and determining a parameter for the aircraft based on a position of the high temperature location about the leading edge of the nose of the aircraft…
However, in the same field of endeavor, Li teaches:
identifying a high temperature location out of the plurality of measurement locations with a highest measured temperature in the temperature distribution (translated document of Li, para. 30: “Low speed - sub-acoustic airfoil angle of attack, including low-speed - sub-acoustic airfoil, airfoil in the air, low speed - sub-acoustic airfoil sensor array layout, low speed - sub-acoustic airfoil leading edge temperature distribution, And the relationship between the angle of attack, in the course of the movement of the aircraft, due to the direction of the flow direction of the thermal element measured film surface temperature Tc will be the largest [i.e., a highest measured temperature in the temperature distribution], the remaining position with the flow of air, Tc will gradually reduce the formation of different angles…”); and
determining a parameter for the aircraft based on a position of the high temperature location about the leading edge of the nose of the aircraft (translated document of Li, para. 30: “Low speed - sub-acoustic airfoil angle of attack, including low-speed - sub-acoustic airfoil, airfoil in the air, low speed - sub-acoustic airfoil sensor array layout, low speed - sub-acoustic airfoil leading edge [i.e., about the leading edge of the nose of the aircraft] temperature distribution, And the relationship between the angle of attack [i.e., determining a parameter], in the course of the movement of the aircraft, due to the direction of the flow direction of the thermal element measured film surface temperature Tc will be the largest [i.e., based on a position of the high temperature location], the remaining position with the flow of air, Tc will gradually reduce the formation of different angles”; Note: To one of ordinary skill in the art, at the time of the application, a leading edge of an airfoil and a leading edge of a nose of an aircraft would be obvious variants for determining aircraft attitudes, see Clark, col. 2, lines 55-59: “An aircraft member, designated generally as 10, includes exposed skin surfaces 12 and 14. While the member 10 as shown is the nose of the aircraft, any other members, such as a curved wing surface, rudder, elevator or other aircraft skin surface member, may be used.” and col. 8, line 61 – col. 9, line 2: “Additional thermal radiation devices having respective fiberoptic cables and photodetectors coupled thereto may be positioned at various locations on the aircraft to determine the status of various aircraft members to aid the pilot and computer in control of the aircraft. For example, thermal radiation members 70, 72, 74 and 76, as shown in FIGS. 8 and 9, positioned along the leading edges of the wing members, both top and bottom, permit the angle of attack of the wings to be precisely determined.”)…
Therefore, it would have been obvious to one of ordinary skill in the art, before the effective filing date of the claimed invention and with a reasonable likelihood of success to modify the invention disclosed by Clark, with the concept of measuring temperature at a plurality of locations about a leading edge of an aircraft component and determine the highest measured temperature, taught by Li, in order to accurately determine an attitude (like the angle of attack) of an aircraft (translated document of Li, para. 50: “…the surface temperature Tc of the thin film layer of the thermal sensor measured at the direction facing the incoming flow will be the largest, while the rest of the positions will gradually decrease with the flow of the airflow, forming different angles of attack. Corresponding to different low speed-subsonic airfoil leading edge temperature distribution, with the change of the angle of attack, different sensitive units of the micro flexible hot film sensor array will measure the maximum value, and the peak value will also move with the change of the angle of attack, as shown in The relationship between the extreme temperature position and the angle of attack; when the angle of attack is 0°, the extreme temperature is at the origin; when the angle of attack is positive, the extreme temperature is at a position on the negative semi-axis; when the angle of attack is negative, the temperature The extreme value is located at a certain position on the positive semi-axis, so that the change in the angle of attack can be converted into the position change of the temperature extreme value along the latitude direction, and the data collected in step 2) can be used to fit the temperature extreme value of the low-subsonic airfoil The relationship curve between the position and the angle of attack can be used to determine the angle of attack by the position of the temperature extreme value measured in real time during the flight…”).
Regarding claim 12, Clark and Li teach The aircraft of claim 11, and Clark further discloses the following:
wherein the parameter is indicative of a relative wind direction during a flight of the aircraft (Clark, col. 3, lines 39-46: “The temperature of the skin of an aircraft flying at high supersonic or hypersonic speeds is proportional to the angle of the skin surface with respect to the flow stream of air [i.e., parameter is indicative of a relative wind direction during a flight of the aircraft] for a given velocity of the aircraft at a given air density. Further, at a given air density and angle of attack, the heating due to air friction is proportional to the velocity of aircraft in the air.”).
Regarding claim 14, Clark and Li teach The aircraft of claim 11, and Clark further discloses the following:
wherein the one or more processors are further configured to execute the instructions stored on the memory to cause the one or more processors to determine an angle of attack of the aircraft based on the parameter (Clark, col. 3, lines 21-26: “Radiation sensors 26 and 28 transmit the electrical data to the aircraft control center, such as computer 30. The data are analyzed in the computer 30 [i.e., one or more processors are further configured to execute the instructions stored on the memory to cause the one or more processors] to determine aircraft flight parameters such as velocity and attitude. The aircraft flight parameters are used by either the computer, the pilot, or both, to control the aircraft.”; col. 3, line 63-col. 4 line 1: “Therefore, any differences in the thermal radiation between the members correspond to differences in angle of the skin with respect to the airflow. For example, assuming θ is 10° and ɑ is zero, then the angle of attack [i.e., determine an angle of attack of the aircraft based on the parameter] of the nose of the plane is zero and the radiation emitted by the respective members 20 and 22 will be equal.”).
Regarding claim 15, Clark and Li teach The aircraft of claim 11, and Clark further discloses the following:
wherein the one or more processors are further configured to execute the instructions stored on the memory to cause the one or more processors to determine a sideslip angle of the aircraft based on the parameter (Clark, col. 3, lines 21-26: “Radiation sensors 26 and 28 transmit the electrical data to the aircraft control center, such as computer 30. The data are analyzed in the computer 30 [i.e., one or more processors are further configured to execute the instructions stored on the memory to cause the one or more processors to] to determine aircraft flight parameters such as velocity and attitude. The aircraft flight parameters are used by either the computer, the pilot, or both, to control the aircraft.”; col. 2, line 62-68: “When member 10 is the nose of the aircraft, the angle of attack, α, illustrated in FIG. 1, is the pitch of the aircraft as the aircraft climbs or dives. Other aircraft attitudes, such as the yaw [i.e., determine a sideslip angle of the aircraft based on the parameter] and roll, may also be measured according to this invention following the same principles used in measuring the pitch.”).
Claim(s) 13 is/are rejected under 35 U.S.C. 103 as being unpatentable over Clark, in view of Li, as applied to claim 11, and further in view of US-20210048003-A1, hereinafter “Egedal” (previously of record).
Regarding claim 13, Clark and Li teach The aircraft of claim 11, but do not appear to explicitly teach the following:
wherein the parameter is indicative of a stagnation point about the leading edge of the nose of the aircraft.
However, in the same field of endeavor, Egedal teaches:
wherein the parameter is indicative of a stagnation point about the leading edge of the nose of the aircraft (Egedal, para. 0015: “The determination of the location of the stagnation point [i.e., a stagnation point about the leading edge of the nose of the aircraft] and/or the determination of the stagnation temperature may be useful for estimating one or more parameters of the air flow characteristics, in particular the angle of attack and the free flow speed.”; Note: One of ordinary skill in the art, at the time of the application, would find temperature characteristics, due to fluid dynamics, of the leading edge of an airfoil of a wind turbine to be an obvious variant to the temperature characteristics, due to fluid dynamics, of the leading edge of an aircraft component.).
Therefore, it would have been obvious to one of ordinary skill in the art, before the effective filing date of the claimed invention and with a reasonable likelihood of success to modify the invention disclosed by Clark, as modified by Li, with the concept of determining a stagnation point along a leading edge of an airfoil or aircraft component, taught by Egedal, in order to accurately determine an attitude of the airfoil or aircraft component, like angle of attack (Egedal, para. 0008: ” Thus, there may be a need for a method and for an arrangement for determining a characteristic of air flow close to a surface of a rotating blade of a wind turbine, which provides a reliable estimation of the flow characteristics and which is easily implemented. Further, there may be a need for a rotor blade for a wind turbine for which a characteristic of air flow close to a surface of the blade can be estimated in an easy manner and in a simple manner, while the blade is rotating.”).
Claim(s) 16-17, and 19 is/are rejected under 35 U.S.C. 103 as being unpatentable over Clark and further in view of Egedal (previously of record) and Islam (previously of record).
Regarding claim 16, Clark discloses A method of operation, comprising:
providing an aircraft comprising a fuselage, a wing, a propulsion system, and a flight information system, the wing connected to the fuselage (Clark, FIG. 8-9: [i.e., an aircraft comprising a fuselage, a wing, a propulsion system…the wing connected to the fuselage]; col. 3, lines 22-27: “Radiation sensors 26 and 28 transmit the electrical data to the aircraft control center [i.e., a flight information system], such as computer 30. The data are analyzed in the computer 30 to determine aircraft flight parameters such as velocity and attitude. The aircraft flight parameters are used by either the computer, the pilot, or both, to control the aircraft.”), and…
…one or more temperature sensors and a controller (Clark, col. 3, lines 22-27: “Radiation sensors 26 and 28 [i.e., one or more temperature sensors] transmit the electrical data to the aircraft control center, such as computer 30 [i.e., a controller]. The data are analyzed in the computer 30 to determine aircraft flight parameters such as velocity and attitude. The aircraft flight parameters are used by either the computer, the pilot, or both, to control the aircraft.”);
determining an aircraft parameter indicative of a pitch attitude of the aircraft with respect to a horizon line during the flight of the aircraft (Clark, col. 4, lines 37-49: “The yaw, roll and other attitudes may be calculated [i.e., determining an aircraft parameter] using appropriately placed thermal radiators and sensors. For each attitude to be sensed, two thermal radiators are placed on either side of the axis of interest. For each aircraft member or aircraft flight parameter of interest, the radiation emitted by two or more sensors appropriately placed can be compared and analyzed to determine the flight status, as has been described with respect to determining the pitch of the nose [i.e., indicative of a pitch attitude of the aircraft with respect to a horizon line during the flight of the aircraft]. For example, to measure the pitch of a member, thermal radiators may be positioned on an upper and lower surface, and to measure yaw of a member, thermal radiators may be positioned on left and right surfaces, etc.”);
determining a wind parameter indicative of an angle of a relative wind flowing in a relative wind direction during the flight of the aircraft (Clark, col. 3, lines 47-55: “The relationship between the angle of attack of the skin surface and the heat flux of the skin surface is illustrated in FIG. 5. FIG. 5 is based on a leading edge of an aircraft having a radius of four inches, a constant air density, and a constant aircraft speed as the angle of the skin surface varies. When the angle of the skin surface with respect to the airflow [i.e., determining a wind parameter indicative of an angle of a relative wind flowing in a relative wind direction during the flight of the aircraft] is between 5° and approximately 20°, the heat flux of the skin surface due to air friction is generally linear, as can be seen from FIG. 5.”; col. 4, lines 50-56: “If the aircraft member 10 does not extend uniformly with respect to a desired plane of measurement, such as in a wing member, the nose of some aircraft, etc., the attitude analysis must take into account differences in angles of the surfaces with respect to each other, differences in air density, and airflow patterns over the different surfaces, and possibly numerous other factors.”)…
…determining an offset angle between the pitch attitude of the aircraft and the relative wind by subtracting the angle of the relative wind from the pitch attitude (Clark, col. 3, line 63-col. 4 line 1: “Therefore, any differences in the thermal radiation between the members correspond to differences in angle of the skin with respect to the airflow [i.e., relative wind direction]. For example, assuming θ is 10° and ɑ [i.e., aircraft parameter] is zero, then the angle of attack [i.e., determining an offset angle between the pitch attitude of the aircraft and the relative wind by subtracting the angle of the relative wind from the pitch attitude] of the nose of the plane is zero and the radiation emitted by the respective members 20 and 22 will be equal.”);
wherein the component of the aircraft is the wing (Clark, col. 2, lines 56-59: “An aircraft member, designated generally as 10, includes exposed skin surfaces 12 and 14. While the member 10 as shown is the nose of the aircraft, any other members, such as a curved wing [i.e., the component of the aircraft is the wing] surface, rudder, elevator or other aircraft skin surface member, may be used.”); and
wherein one or more temperature sensors are disposed within an interior of the wing outside of an external environment of the aircraft (Clark, FIGs. 1 and 4; col. 2, lines 56-59: “While the member 10 as shown is the nose of the aircraft, any other members, such as a curved wing surface [i.e., the wing], rudder, elevator or other aircraft skin surface member, may be used.”; col. 3, lines 4-7: “Fiberoptic cables 24 are mounted directly adjacent the thermal radiation members 20 and 22 for guiding the thermal radiation to respective radiation detectors 26 and 28 [i.e., one or more temperature sensors].”; col. 5, line 63 - col. 6, line 1: “Alternatively, as shown in FIG. 4, the thermal radiation member 20 may be mounted directly behind skin surface 12. Heat is transferred by skin surface 12 to thermal radiation member 20. The thermal radiation member 20 of the embodiment of FIG. 4 is mounted in direct contact with the skin member 12, but not in contact with the external atmosphere [i.e., disposed within an interior of the wing outside of an external environment of the aircraft].”).
Clark does not appear to explicitly disclose the following:
…the flight information system comprising a gyroscope onboard the aircraft…the wind parameter determined based on a position of a stagnation point about a leading edge of a component of the aircraft, the position of the stagnation point being based on a plurality of temperatures of the leading edge of the component of the aircraft measured by the one or more temperature sensors; and…
However, in the same field of endeavor, Egedal teaches:
…the wind parameter determined based on a position of a stagnation point about a leading edge of a component of the aircraft (Egedal, para. 0015: “The determination of the location of the stagnation point [i.e., based on a position of a stagnation point] and/or the determination of the stagnation temperature may be useful for estimating one or more parameters of the air flow characteristics [i.e., determining wind parameter], in particular the angle of attack and the free flow speed.”),
the position of the stagnation point being based on a plurality of temperatures of the leading edge of the component of the aircraft measured by the one or more temperature sensors (Egedal, para. 0012-0013: “For measuring at least one value of the temperature of air close to the surface of the blade, one or more temperature sensors [i.e., measured by the one or more temperature sensors] installed in or at the blade, in particular such as that the temperature of the air close to the surface of the blade is [measurable], may be utilized. The temperature sensor(s) may be a conventional temperature sensor(s). The at least one value of the temperature may in particular be measured close to (an expected location of) a stagnation point or a location of a stagnation point may be determined or estimated using at least one value or plural values of temperature of the air close to the surface of the blade. In particular, a temperature field at different locations of the surface of the rotor blade may be measured and based on the temperature field or based on plural temperature values, a location of a stagnation point may be determined [i.e., position of the stagnation point being based on a plurality of temperatures of the leading edge of the component of the aircraft measured by the one or more temperature sensors] and a stagnation temperature may be determined.”; para. 0014: “For high-speed flows (transonic and supersonic, i.e. with speeds close to or higher than the speed of sound) it may be well-known that there is an increase of temperature of the flow in the stagnation points. The stagnation points of the flow are those points (or regions) at which the flow (substantially) comes to rest (e.g. having zero velocity), e.g. a location close to the leading edge of an airfoil [i.e., temperatures of the leading edge of the component] of a rotor blade.”; Note: One of ordinary skill in the art, at the time of the application, would find temperature characteristics, due to fluid dynamics, of the leading edge of an airfoil of a wind turbine to be an obvious variant to the temperature characteristics, due to fluid dynamics, of the leading edge of an aircraft component.);
Therefore, it would have been obvious to one of ordinary skill in the art, before the effective filing date of the claimed invention and with a reasonable likelihood of success to modify the invention disclosed by Clark, with the concept of determining the location of a stagnation point on the leading edge of an airfoil or a component of an aircraft to determine a wind parameter, taught by Egedal, in order to accurately estimate parameters related to air flow and airfoil or aircraft orientation (Egedal, para. 0029: “The location of the stagnation point is expected to be close to the leading edge. Thus, measuring the temperature values at different positions around the leading edge may enable to accurately determine the location of the stagnation point which may in turn enable for determining the direction of the free air flow in a reliable and accurate manner.”).
Clark and Egedal do not appear to explicitly disclose the following:
…the flight information system comprising a gyroscope onboard the aircraft…
However, in the same field of endeavor, Islam teaches:
…the flight information system comprising a gyroscope onboard the aircraft (Islam, Abstract: “AHRS is an important system that provides the 3-dimensional orientation (roll, pitch and yaw or heading) of an aircraft from which the flying performance can be evaluated.”; col. 1, last paragraph: “An IMU generally comprises of three axis gyroscopes, accelerometers and magnetometers for measuring the 3D orientation (roll, pitch and yaw angle).”; col. 2, paragraph 3: “This paper proposes a standalone AHRS utilizing low cost MEMS based three axis accelerometers, gyroscopes and magnetometers using a complementary filter to determine the flying performance of a low speed aircraft.”)…
Therefore, it would have been obvious to one of ordinary skill in the art, before the effective filing date of the claimed invention and with a reasonable likelihood of success to modify the invention disclosed by Clark, as modified by Egedal, with the concept of using at least one gyroscope to determine orientation of an aircraft, taught by Islam, in order to maintain control of the aircraft based on accurate estimates of speed and orientation (Islam, Abstract: “AHRS is an important system that provides the 3-dimensional orientation (roll, pitch and yaw or heading) of an aircraft from which the flying performance can be evaluated.”).
Regarding claim 17, Clark, Egedal, and Islam teach The method of claim 16, and Clark further discloses the following:
wherein the determining of the wind parameter comprises measuring a temperature at a plurality of measurement locations about the leading edge of the component of the aircraft to determine a temperature distribution about the leading edge of the component of the aircraft (Clark, col. 3, lines 7-10: ”The quantity of thermal radiation emitted to the sensors 26 and 28 is detected, and an electrical signal is generated corresponding to the quantity of thermal radiation received [i.e., comprises measuring a temperature at each of a plurality of measurement locations about the leading edge of the component of the aircraft to determine the temperature distribution].”; col. 3, lines 47-55: “The relationship between the angle of attack of the skin surface and the heat flux of the skin surface is illustrated in FIG. 5. FIG. 5 is based on a leading edge [i.e., a plurality of measurement locations about the leading edge of the component of the aircraft] of an aircraft having a radius of four inches, a constant air density, and a constant aircraft speed as the angle of the skin surface varies. When the angle of the skin surface with respect to the airflow is between 5° and approximately 20°, the heat flux of the skin surface due to air friction is generally linear, as can be seen from FIG. 5.”; col. 4, lines 50-56: “If the aircraft member 10 does not extend uniformly with respect to a desired plane of measurement, such as in a wing member, the nose of some aircraft, etc., the attitude analysis [i.e., determining a wind parameter] must take into account differences in angles of the surfaces with respect to each other, differences in air density, and airflow patterns over the different surfaces, and possibly numerous other factors.”).
Clark does not appear to explicitly disclose the following:
identifying a high temperature location out of the plurality of measurement locations with a highest measured temperature in the temperature distribution; and determining the position of the stagnation point based on a position of the high temperature location about the leading edge of the component of the aircraft.
However, Egedal further teaches the following:
identifying a high temperature location out of the plurality of measurement locations with a highest measured temperature in the temperature distribution; and determining the position of the stagnation point based on a position of the high temperature location about the leading edge of the component of the aircraft (Egedal, para. 0014: “For high-speed flows (transonic and supersonic, i.e. with speeds close to or higher than the speed of sound) it may be well-known that there is an increase of temperature of the flow in the stagnation points. The stagnation points of the flow are those points (or regions) at which the flow (substantially) comes to rest (e.g. having zero velocity), e.g. a location close to the leading edge of an airfoil of a rotor blade. The increase of the temperature at the stagnation point [i.e., determining the position of the stagnation point based on a position of the high temperature location about the leading edge of the component] may be understood based on the first law of thermodynamics.”; para. 0058: “Thus, the increase in the temperature is proportional to the square of the free flow velocity. The increase in temperature is highest at the stagnation point (i.e. where the flow comes to rest and all kinetic energy is transformed to thermal energy/pressure) [i.e., identifying a high temperature location out of the plurality of measurement locations with a highest measured temperature in the temperature distribution].”).
Therefore, it would have been obvious to one of ordinary skill in the art, before the effective filing date of the claimed invention and with a reasonable likelihood of success to modify the invention disclosed by Clark, as modified by Egedal and Islam, with the concept of determining the location of a stagnation point on the leading edge of an airfoil or a component of an aircraft to determine a wind parameter, taught by Egedal, in order to accurately estimate parameters related to air flow and airfoil or aircraft orientation (Egedal, para. 0029: “The location of the stagnation point is expected to be close to the leading edge. Thus, measuring the temperature values at different positions around the leading edge may enable to accurately determine the location of the stagnation point which may in turn enable for determining the direction of the free air flow in a reliable and accurate manner.”).
Regarding claim 19, Clark, Egedal, and Islam teach The method of claim 16, and Clark further discloses the following:
…the offset angle comprises an angle of attack or a sideslip angle of the aircraft (Clark, col. 3, lines 58-61: “The angle of attack [i.e., offset angle comprises an angle of attack] can be determined by measuring the difference between the thermal radiation emitted by members 20 and 22.”; col. 2, lines 65: “Other aircraft attitudes, such as the yaw [i.e., or a sideslip angle of the aircraft] and roll, may also be measured according to this invention following the same principles used in measuring the pitch.”).
Clark and Egedal do not appear to explicitly disclose the following:
wherein the aircraft parameter is determined using the gyroscope onboard the aircraft; and…
However, in the same field of endeavor, Islam teaches:
wherein the aircraft parameter is determined using a gyroscope onboard the aircraft (Islam, Abstract: “AHRS is an important system that provides the 3-dimensional orientation (roll, pitch and yaw or heading) of an aircraft from which the flying performance can be evaluated.”; col. 1, last paragraph: “An IMU generally comprises of three axis gyroscopes, accelerometers and magnetometers for measuring the 3D orientation (roll, pitch and yaw angle).”; col. 2, paragraph 3: “This paper proposes a standalone AHRS utilizing low cost MEMS based three axis accelerometers, gyroscopes and magnetometers using a complementary filter to determine the flying performance of a low speed aircraft.”); and…
Therefore, it would have been obvious to one of ordinary skill in the art, before the effective filing date of the claimed invention and with a reasonable likelihood of success to modify the invention disclosed by Clark, as modified by Egedal and Islam, with the concept of using a gyroscope to determine an aircraft parameter, taught by Islam, in order to maintain control of the aircraft based on accurate estimates of speed and orientation (Islam, Abstract: “AHRS is an important system that provides the 3-dimensional orientation (roll, pitch and yaw or heading) of an aircraft from which the flying performance can be evaluated.”).
Conclusion
Applicant's amendment necessitated the new ground(s) of rejection presented in this Office action. Accordingly, THIS ACTION IS MADE FINAL. See MPEP § 706.07(a). Applicant is reminded of the extension of time policy as set forth in 37 CFR 1.136(a).
A shortened statutory period for reply to this final action is set to expire THREE MONTHS from the mailing date of this action. In the event a first reply is filed within TWO MONTHS of the mailing date of this final action and the advisory action is not mailed until after the end of the THREE-MONTH shortened statutory period, then the shortened statutory period will expire on the date the advisory action is mailed, and any nonprovisional extension fee (37 CFR 1.17(a)) pursuant to 37 CFR 1.136(a) will be calculated from the mailing date of the advisory action. In no event, however, will the statutory period for reply expire later than SIX MONTHS from the mailing date of this final action.
Any inquiry concerning this communication or earlier communications from the examiner should be directed to Leah N Miller whose telephone number is (703)756-1933. The examiner can normally be reached M-Th 8:30am - 5:30pm ET.
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/L.N.M./Examiner, Art Unit 3666
/HELAL A ALGAHAIM/SPE , Art Unit 3666