The origin of the ONSPEED Aural AOA logic is the system originally installed in the McDonnell F-4 Phantom. These are excerpts from the "Dash One" (flight manual) for Double Ugly, as the F-4 was affectionately known by those of us that flew the airplane. All US fighter aircraft use AOA as a primary reference for approach and landing and have cockpit cuing systems (visual and aural) that assist the pilot during approach and landing. Cues vary from airplane to airplane, but most US fighters use the conventional doughnut/chevron visual display in addition to other HUD or instrument-based energy cues. These old flight manual excerpts are representative of AOA and energy management techniques employed by fighter pilots in the traffic pattern, including coming aboard ship.
Summary: The rate of general aviation (GA) accidents and fatalities is the highest of all aviation categories and has been nearly constant for the past decade. According to the National Transportation Safety Board, in 2010, GA accidents accounted for 96% of all aviation accidents, 97% of fatal aviation accidents, and 96% of all fatalities for U.S. civil aviation. However, GA accounted for 51% of the estimated total flight time of all U.S. civil aviation in 2010.
The FAA identified the top three causes of fatal GA accidents to be: 1) loss of control (LOC) in flight, 2) controlled flight into terrain, and 3) system or component failure/power plant. Additionally, the General Aviation Joint Steering Committee (GAJSC) recently published its final report on LOC, approach, and landing. The GAJSC LOC working group recommended angle of attack (AoA) systems as one of its top safety enhancements for GA aircraft. Current differential pressure AoA systems for light GA aircraft concentrate on the slow flight/stall regime. Typically,outside of this flight regime, these AoA systems are inaccurate. Therefore, they are not usable forcritical flight regimes other than the stall region. This study examined such systems. It was found that using unnormalized differential pressure (Pfwd-P45) does not provide adequate accuracy throughout the aircraft’s AoA range. Using unnormalized differential pressure (Pfwd-P45) does not yield accurate AoA data throughout the aircraft’s normal operating envelope. The calibration curve is nonlinear. For a limited range of high AoA near stall, a linear fit to the data near stall provides adequate accuracy. However, accuracy at low AoA, such as that required by cruise, is poor. Therefore, systems using unnormalized differential pressure—similar to that tested, which use a linear calibration—are basically stall warning devices.Four alternate techniques were flight tested using the two pressure ports designated as Pfwd and P45on the probe used with the commercial off-the-shelf (COTS) AoA data acquisition system (DAS). The flight test configurations were: the ratio of Pfwd/P45; (Pfwd-P45)/P45, which is just (Pfwd/P45)-1; (Pfwd-P45)/q; and (Pfwd-P45). Only the ratio of Pfwd/P45 provided an accurate AoA throughout the aircraft’s normal operating environment, including into and recovery from the stall region. The calibration curve based on the ratio of Pfwd/P45 was studied and determined to be linear throughout the aircraft’s AoA provided that the probe was located on the wing’s lower surface between an estimated 25% and 60% of the local wing chord. The resulting calibration curve was linear because of the smooth laminar flow on the lower wing surface including when the upper wing surface was partially or fully separated. Normalizing (Pfwd-P45)/q (i.e., normalizing differential pressure with dynamic pressure) also produced a linear calibration curve, provided that the aircraft’s true freestream dynamic pressure was used. Similar results may be expected for other differential pressure systems. A low-cost ($100/table A-1 of appendix A) differential, pressure-based COTS AoA DAS was designed, successfully reduced to practice, wind tunnel tested, and flight tested. The accuracy of the COTS differential pressure AoA system was determined to be 1 ⁄ 4 to 1 ⁄ 2 of a degree. The repeatability of the data from the COTS system was excellent. Differential pressure AoA systems are dynamic pressure-dependent. A physics-based determination of AoA was successful, provided that a reasonably accurate aircraft lift curve was determined. Calculation of the lift curve slope was within 0.01 degrees of the value determined by the flight test, using an alpha/beta probe.
Summary: This research was conducted as the result of a concern within the FAA’s Small Airplane Directorate regarding AIR Policy Memo AIR100-14-110-PM01, “Approval of Non-Required Angle of Attack (AoA) Indicator Systems.” This policy memo provides relief from typical design approvals to manufacturers of non-required/supplemental angle-of-attack systems. The purpose of the research conducted was to evaluate all of the commercial off-the-shelf angle-of-attack systems available. Each system was installed on a test aircraft, along with an angle-of-attack truth source, and flown through various maneuvers designed to evaluate the ability of the system to accurately and repeatedly measure angle of attack. Test data were analyzed, and any observed deficiencies were reported. It was found that some commercially available angle-of-attack systems, manufactured and installed under AIR100-14-110-PM01, have characteristics that result in false indications of positive stall margin. However, the majority provided safe and appropriate indications throughout the conditions tested. Flight tests and evaluations also considered human factors, aspects of the system’s display, and associated interface. This evaluation also included review of each system’s calibration procedure. Recommendations are provided for manufacturers and FAA/American Society for Testing and Materials (ASTM) to consider for improving the performance of non-required/supplemental angle-of-attack systems.
This memorandum establishes requirements and procedures for issuing a design and production approval to a United States (U.S.) manufacturer under
Title 14 of the Code of Federal Regulations (14 CFR) 21.8(d) for a non-required/supplemental Angle of Attack (AoA) indicator system.
Abstract: A wind tunnel calibration of a commercially available general aviation differential pressure angle of attack probe was conducted. Differential pressure varied linearly with pitch angle for each dynamic pressure tested. Beyond ±6 degrees of yaw angle, differential pressure rolled off rapidly with yaw angle. Normalizing differential pressure with dynamic pressure collapsed the differential pressure to a single curve for pitch. Similarly, normalizing differential pressure with the dynamic pressure collapsed the differential pressure to a single curve for yaw. Normalizing the differential pressure with dynamic pressure removes the effect of speed and density altitude when deriving angle of attack from differential pressure. Normalizing the differential pressure using the pressure from the 45 degree probe surface similarly collapsed the differential pressure to single separate curves for both pitch and yaw. Differential pressure varied parabolically with pitch angle when normalized using the pressure from the 45◦ probe surface. For pitch angles above six degrees, a linear relation between differential pressure normalized with the pressure from the 45 degree probe surface and the square of the pitch angle provided an adequate approximation. The typical two-point linear approximation based on two in-flight calibration points results in significant errors in displayed angle of attack. Moving the low angle of attack in-flight calibration point closer to the stall angle of attack reduces the displayed angle of attack error in the critical stall region. Using a four-point linear approximation significantly reduces the error in displayed angle of attack throughout the angle of attack range. Combined pitch, yaw and roll results in sideslip, which produces significant error
in displayed angle of attack based on calibration at zero yaw and roll.
Abstract: Recommendations resulting from the study of true angle of attack and yaw measurement are presented. The desired accuracy is 0.10 degree under normal flight conditions at Mach numbs up to 3.0 and altitudes up to 100,000 feet. Sensor location on the airframe, the actual sensing device, and the associated computing mechanism are deemed to be the problems of prime importance. A wide variety of methods for determining angle of attack were analyzed with emphasis on sensor sensitivity, computer complexity, and system reliability. Studies were carried through the stage of devising mechanization for systems showing any promise. In general, pressure sensing and lift mechanization methods appear to off the most promise.
Abstract: The Research and Development Division of the United States Air Force Instrument Flight Center conducted an evaluation and study of the Bendix Standardized Angle of Attack (AOA) System for its use in Air Training Command flying training programs. The Bendix AOA system was installed in T-38, SIN 70-1549 for engineering flight test at the Air Force Flight Test Center, Edwards AFB, California. The aircraft was then flown to Randolph AFB for pilot factors evaluation and determination of exactly what flying maneuvers could be flown using AOA as the controlling parameter and how AOA should be used in these maneuvers. Subjective data on the procedures and techniques for AOA use were gathered from twelve T-38 instructor pilots from
the Pilot Instructor Training (PIT) and Instrument Pilot Instructor School (IPIS) at Randolph AFB.
Angle of attack (AOA) is an aerodynamic parameter that is key to understanding the limits of airplane performance. Recent accidents and incidents have resulted in new flight crew training programs, which in turn have raised interest in AOA in commercial aviation. Awareness of AOA is vitally important as the airplane nears stall. It is less useful to the flight crew in the normal operational range. On most Boeing models currently in production, AOA information is presented in several ways: stick shaker, airspeed tape, and pitch limit indicator. Boeing has also developed a dedicated AOA indicator integral to the flight crew’s primary flight displays.
Abstract: Angle of Attack (AOA) is an important aeronautical concept used to understand the performance status of an aircraft during different flight stages. The Federal Aviation Administration (FAA) has indicated the importance of developing and encouraging the use of affordable AOA based systems to increase inflight safety. Embry-Riddle Aeronautical University’s flight department decided to install AOA indicators in its fleet of Cessna 172S, to increase safety and to help student pilots better understand this important concept. This paper presents a review of AOA, visual display design principles, and usability. This experimental study examined three different AOA indicators provided by the flight department. The goal was to conduct a usability study in order to understand which of these indicators was better suited for student training. Instructor pilots were used as participants in a series of flights, in which they were asked to perform different maneuvers in which using AOA indicators was thought to help increasing stall awareness and performance. At the end of each flight participants were asked to complete a series of surveys (including an adaptation of the system usability scale) and to provide comments in order to understand their preferences related to AOA indicators. The analysis of the data shows significant differences between the indicators. Discussion of the results and recommendations for future studies are also covered.
Abstract: The purpose of this study is to explore the use of an Inertial Navigation System as a primary method for measuring aircraft air flow angles in flight testing. The traditional methods used to measure air flow angles consist of sensors external to the aircraft, such as an air data boom or an angle of attack probe. The advantage of using INS to measure air flow angles would be in the simplicity of the instrumentation. All components could be fixed internally, leaving minimal external modifications to the aircraft necessary for instrumentation. This would reduce costs and instrumentation time and enable air flow angle data collection in the many aircraft already fitted with an INS. Other downfalls to external sensors are the complicated calibrations and error corrections that must be used to compensate for upwash and position error of the instruments. This study will use flight test data from the Diamond DA42 Twinstar flight test program, conducted by Embry Riddle Aeronautical University. A method was developed to estimate the air flow angles using INS and other standard flight test parameters that exclude an external air data boom. This method involves determining wind velocity in order to compute an estimate for the air flow angles. Multiple Kalman Filters use air flow angle estimates to determine essential aircraft stability derivatives. Initial values for these stability derivatives are inaccurate but, over a short period of time, the Kalman Filters are able to converge to an accurate solution, provided the necessary parameters are made observable by aircraft dynamics. The converged stability derivatives are combined with aircraft accelerations to produce accurate air flow angle measurements. These air flow angles are validated against the traditionally measured air flow angles. This enables derivation of an error prediction method for INS air flow angle measurements. The predicted error is initially high, but converges along with the estimate of the stability derivatives. The methods developed in this study are implemented in a way such that real-time estimation of the air flow angles would be possible. This method is unique by focusing on instantaneous acceleration measurements while simultaneously estimating stability derivatives.
Abstract: The crucial relation of angle of attack to aircraft performance suggests than an angle of attack instrument Would enhance the process of learning to pilot an airplane. Therefore, a project to determine the possible value of angle of attack presentation in addition to other required instruments for flight training in general aviation aircraft was conducted. The project entailed comparing the performance of two similar groups of Embry-Riddle Aeronautical Institute flight students enrolled in the private pilot course. Flight instruction of both groups proceeded concurrently utilizing the same aircraft except the experimental group was trained using an angle of attack instrument in addition to the airspeed indicator. A series of three scored tests was employed to measure the performance of each student on selected maneuvers during and upon completion of the course. Scores of the experimental group and the control group were tested for significance of difference by the analysis of variance method. A comparison of the derived variance ratios with the corresponding values in the Table of F ratios at the 5% level signified in all instances that the null hypothesis should not be rejected. Consequently, statistical evidence indicated that there was no true difference in the quality of performance of students trained with and without the angle of attack
indicator at the private pilot level. The overall similarity of the performance of the two groups is attributed to the following two conditions. (1) Experimental group students were required to learn the use of the angle of attack indicator in addition to the airspeed indicator. The difficulty certain students experienced early in the program in developing skill in using this instrument tended to compensate for possible enhancing effect whichmight have been realized in the final stage. (2) At the present state of the development of flight instruction curricula, contact flight is the quintessence of
the private pilot program. An instrument capable of producing a significant effect on pilot performance at this level, consequently, would be rare.
Findings of this project indicate thct further research in the use of the angle of attack indicator is appropriate. Projects should be conducted to determine the value of angle of attack presentation: (1) when used in lieu of airspeed in private pilot training, and (2) in instrument flight training.
Abstract: The use of angle-of-attack information for a pilot's display in a general-aviation airplane was investigated to determine whether this form
of information would improve performance and flight safety. An angle-of-unit, and a display instrument was installed and flight tested in a typical
attack system consisting of a wing-mounted vane, an electronic computer twin-engine, general-aviation airplane, The flight-test maneuvers were
limited to the low-speed flight region where the benefits of angle-of-attack presentation were likely to be greatest. Some of the expected advantages
of this parameter, such as visual indication of stall margin and its independence of gross weight and flap position, were realized: however.
certain aerodynamic characteristics of the airplane. such as the phugoid and directional-control capability, were found to limit and tended to
negate some of the expected advantages. As a result this use of angle of attack did not show a significant improvement in performance and
Abstract: The National Aeronautics and Space Administration (NASA) conducted a literature review to determine the potential benefits of a display of angle-of-attack (AoA) on the flight deck of commercial transport that may aid a pilot in energy state awareness, upset recovery, and/or diagnosis of air data system failure. This literature review encompassed an exhaustive list of references available and includes studies on the benefits of displaying AoA information during all phases of flight. It also contains information and descriptions about various AoA indicators such as dial, vertical and horizontal types as well as AoA displays on the primary flight display and the head up display. Any training given on the use of an AoA indicator during the research studies or experiments is also included for review.
Abstract: Experimental tests have been completed which recorded the ability of two combination steady state and high response time varying Pitot probe designs to accurately measure steady stagnation pressure at a single location in a flow field. Tests were conducted of double-barreled and coannular Prati probes in a 3.5 in. diameter free jet probe calibration facility from Mach 0.1 to 0.9. Geometric symmetry and pitch (–40° to 40° ) and yaw (0° to 40° ) angle actuation were used to fully evaluate the probes. These tests revealed that the double-barreled configuration induced error in its steady state measurement at zero incidence that increased consistently with jet Mach number to 1.1 percent at Mach 0.9. For all Mach numbers, the double-barreled probe nulled at a pitch angle of approximately 7.0° and provided inconsistent measurements when yawed. The double-barreled probe provided adequate measurements via both its steady state and high response tubes (within ±0.15 percent accuracy) over unacceptable ranges of biased pitch and inconsistent yaw angles which varied with Mach number. By comparison, the coannular probe provided accurate measurements (at zero incidence) for all jet Mach numbers as well as over a flow angularity range which varied from ±26.0° at Mach 0.3 to ± 14.0° at Mach 0.9. Based on these results, the Prati probe is established as the preferred design. Further experimental tests are recommended to document the frequency response characteristics of the Prati probe.
Abstract. This project determined the position errors of an aircraft's AOA sensor using state estimation with flight test data. The position errors were caused by local flow and upwash and were found to be a function of AOA and Mach number. The test aircraft used in the is project was a T-38A from the USAF Test Pilot School. The position errors were found by calculating the true AOA using equations of motion and DAS parameters. The estimated AOA was compared to the measured AOA from the notebook sensor to obtain the position error. Accurate position errors were obtained, even in dynamic maneuvers. This method should be considered in future AOA error testing.
Abstract: Technically inclined flyers have long recognized the importance to the pilot of angle of attack information. Unfortunately, a reliable, accurate, and low cost angle of attack indicator has not been available. From the outset, the goal of this research has been to develop an angle of attack indicator as simple, accurate, reliable, and inexpensive as the common airspeed indicator so that it could be used as a primary landing approach aid and for the other
flight control functions. This report describes the solution to that problem. The pressure driven angle of attack indicating system usually consists of an external probe and an internally mounted indicator, the two elements being connected by three tubes. The approach is mainly experimental although
extensive analytics are provided to aid engineers in the design of pressure driven angle of attack indicating systems for almost any application.
Accuracy of -0.08 degree was demonstrated in the wind tunnel. Accuracy of 0.17 degree was demonstrated in flight. Flight tests showed the response
characteristics to be ideal for pilot use. The pressure driven angle of attack indicating system equals or exceeds the best of other systems in all
Abstract: Throughout aviation history, pilots and engineers have had to rely on mechanical angle of attack and sideslip probes to determine an aircraft's position relative to the airmass. Recent advances in the stability, accuracy and reliability of inertial navigation and reference systems now allow angle of attack and sideslip information to be calculated from internal aircraft systems and a central air data computer. Conflicting requirements for inflight angle
of attack information and post-flight angle of attack and sideslip data reduction demand two separate methods. In flight algorithms require fast, accurate angle of attack, with no assumptions on vertical wind. Post-flight usage, however, demands great accuracy with no assumptions on either sideslip or vertical windage. From the aircraft equations of motion, angle of attack and sideslip algorithms will be developed, with velocity and rate inputs of the type expected from an aircraft central air data computer and inertial navigation system. A computer program will then be developed to validate these equations. A Kalman filter algorithm will also be designed to aid in estimating data output from these sources.
Abstract: The purpose of this report is to provide system development personnel with a set of general guidelines for evaluating a newly developed cockpit alerting and warning system in terms of human factors issues. Although the discussion centers around a general methodology, it has been made
specific to the issues involved in alerting systems. The approach has been to look to the future in preparation for next generation commercial aircraft and the application of a mole mature technology of automation. An overall statement of the current operational problem is presented, with an attempt to describe the more salient human factors problems with reference to existing alerting and warning systems. Next, the methodology for proceeding through system development to system test is discussed, with special emphasis on the differences between traditional human factors laboratory evaluations and those required for evaluation of complex man-machine systems under development. The last section deals more explicitly with performance evaluation in the alerting and warning subsystem using a hypothetical sample system, A further implicit purpose of this report is to engender an industry consensus as to a logical, efficient, and economical way to proceed to a new generation solution of the alerting system problem.
Abstract: The NASA Dryden Flight Research Facility embarked upon a project with the United States Army Aviation Engineering Flight Activity (USAAEFA) to develop and test a stall-speed warning system. NASA designed and built an automated stall-speed warning system which presents both airspeed and stall speed to the pilot. The airspeed and stall speed are computed in real time by monitoring the basic aerodynamic parameters (dynamic pressure, horizontal and vertical accelerations, and pressure altitude) and other parameters (elevator and flap positions, engine torques, and fuel flow). In addition, an aural warning at predetermined stall margins is presented to the pilot through a voice synthesizer. Once the system was designed and installed in the aircraft, a flight-test program of less than 20 hr was anticipated to determine the stall speed software coefficients. These coefficients would then be inserted in the system's software and then test flown over a period of about 10 hr for the purposes of evaluation.
Abstract: Implementation of the Next Generation Air Transportation System (NextGen) will require shifting more roles to the flight deck. The proposed tools and displays for facilitating these added tasks primarily deliver information through visual means. This saturates an already loaded channel while perhaps underutilizing the auditory modality. This paper describes audio enhancements we have developed to compliment NextGen tools and displays, and reports on preliminary observations from a simulation incorporating these enhancements. Pilots were generally receptive to the broad concept, but opinions diverged regarding specific features, suggesting potential for this strategy, and that user defined settings may be important.
Abstract: The goal of this research is increased safety in aviation. Aviation is a highly automated and complex, as well as, safety critical human-machine system. The pilot communicates with the system via a human-machine interface in cockpit. In an alerting situation this interface is in part an auditory alerting system. Human errors are often consequences of actions brought about by poor design. Pilots complain that they may be both disturbed and annoyed of alerts, which may effect performance, especially in non-normal situations when the mental workload is high. This research is based on theories in ergonomics and cognitive engineering with the assumption that improved human performance within a system increase safety. Cognitive engineering is a design philosophy for reducing the effort required by cognitive functions by changing the technical interface, which may lead to improved performance. Knowledge of human abilities and limitations and multidisciplinary interrelated theories between humans, sounds and warnings are used. Several methods are involved in this research, such as literature studies, field studies, controlled experiments and simulations with pilots. This research defines design requirements for sounds appropriate in auditory alerts as Natural Warning Sounds. For example, they have a natural meaning within the user’s context, are compatible with the auditory information process, are pleasant to listen to (not annoying), are easy to learn and are clearly audible. A design process for auditory alerting systems is suggested. It includes methods of associability and sound imagery, which develop Natural Warning Sounds, and combines these with an appropriate presentation format. Associability is introduced and represents the required effort to associate sounds
to their assigned alert function meaning. An associable sound requires less effort and fewer cognitive resources. Sound imagary is used to develop soundimages. A sound image is a sound, which by its acoustics characteristics has a particular meaning to someone without prior training in a certain context. Simulations of presentation formats resulted in recommendations for cancellation capabilities and avoiding continuously repeated alerts.
This research brings related theories closer to practice and demonstrates general methods that will allow designers, together with the users of the system, to apply them in their own system.
Abstract: The only way to simplify and promote the effective use of an alerting system that must be comprehensive in its coverage of hazardous or non-normal conditions is to convey top level information that provides an indication of criticality and identity. In an attempt to reduce the number of aural alerting signals presented in aircraft flight decks, this investigation pursued advances toward the development of a simple aural alert categorization scheme that provides flight deck function and urgency level information. In Experiment 1, 20 subjects having “normal” hearing threshold levels provided magnitude estimation urgency ratings for a series of aural alerts. These ratings revealed that subjects perceived low, moderate, and high urgency levels within each of four equally urgent aural alerting sets. In Experiment 2, 12 subjects having “normal” hearing threshold levels participated in a brief training session and then performed a sound identification task in conjunction with an automated and manual tracking task. Sound identification data revealed that subjects correctly identified the alerting set (i.e., major flight deck function) and urgency level associated with each of 12 aural alerts in 96.53% of the trials occurring during automated tracking and in 95.83% of the trials occurring during manual tracking; furthermore, subjects correctly identified each alerting set, urgency level, and aural alert equally often during each tracking task condition. Electroencephalogram (EEG) data recorded throughout the
performance of each tracking task condition revealed that manual tracking required a significantly higher level of attentional engagement than automated tracking. Subjective assessments of workload collected after the performance of each tracking task condition revealed that a significantly higher level of workload was experienced during the manual condition of the tracking task than during the automated condition of the tracking task. Collectively, this investigation’s results indicated that acoustic parameter manipulations can be used to create four distinctive alerting sets that each convey three levels of urgency and that these alerting sets and urgency levels can be accurately identified when two levels of workload and attentional engagement are
Abstract: Engineers and pilots rely on mechanical flow angle vanes on air data probes to determine the angle of attack of the aircraft in flight. These probes, however, are costly, come with inherent measurement errors, affect the flight characteristics of the aircraft, and are potentially dangerous in envelope expansion flights. Advances in the accuracy, usability, and affordability of inertial navigation systems allow for angle of attack to be determined accurately without direct measurement of the airflow around the aircraft. Utilizing an algorithm developed from aircraft equations of motion, a post- flight data review is completed as the first step in proving the low cost feasibility of utilizing inertial navigation data for such analysis. Flight tests were conducted with the UTSI Cessna 210 research aircraft to calibrate an angle of attack flow angle vane and obtain inertial navigation data from a commercial INS system in typical flight scenarios. The results of the angle of attack algorithm are compared to the measured angle of attack flow angle vane. Discussed in this thesis are the feasibility and potential applications of angle of attack determination from inertial data.
Abstract: This technical information memorandum documents a theoretically-based method of estimating the upwash angle at nose boom-mounted vanes of an arbitrary aircraft. The method is commonly called the Yaggy-Rogallo technique and has been used successfully for many years at the NASA/Dryden Flight Research Center and the Air Force Flight Test Center. The technique divides any aircraft into a collection of bodies of revolution and thin airfoils, makes upwash angle estimates for each component, and combines at the individual estimates into a single estimate. Development of the theory for upwash angle estimates of both bodies of revolution and thin airfoils are presented. Recent improvements in solution and use of the Yaggy-Rogallo equations are also presented. Guidelines for using the upwash angle estimates in the data reduction an analysis are presented along with comparisons to flight test upwash calibration data showing excellent results. FORTRAN V computer programs were written so simplify and expedite making upwash angle estimates. Listings of the program code, a user's guid and a programmer's guide are included in the memorandum.
Abstract: A summary of equations often used in free-flight and wind-tunnel data reduction and analysis is presented. Included are transfer equations for accelerometer, rate gyro,and angle-of-attack instrumentation; axesysystem transfers fo aerodynamic derivatives; and methods for measuring moments of inertia. In general, the equations are in a complete form; for example, those terms are retained that are missing when planar symmetry is assumed for airplanes.
Abstract. This report shows how an aircraft’s energy state and energy rate capabilities are directly related to operational maneuverability and efficiency in terms of energy-maneuverability theory. It demonstrates also how EM theory may be applied to assist the tactician, commander, planner and designer in optimizing aircraft performance. Load factor verses velocity (G-V) and altitude vs Mach number (H-M) diagrams are employed to obtain the interacting energy relationships fundamental to EM theory. The C-V diagrams provide a measure of instantaneous maneuverability with the H-M diagrams (the most valuable diagrams) show sustained maneuverability as a function of energy rate, G, efficiency and range throughout an aircraft’s performance envelope. The energy diagrams as the working tools of EM theory, may be used to determine operation maneuverability and efficiency of various armament-engine-airframe combinations.
Abstract. This white paper details the FlyONSPEED aural logic and hardware. Flight test data as of April 2021 are summarized. Includes detailed description of system operation, theory, project history and flight test results.
10 Slide PowerPoint briefing detailing the FlyONSPEED project. Includes 2 embedded videos, 216 MB file size.
The effect of rate of change of angle of attack on the maximum lift coefficient of a pursuit airplane equipped with a low-drag-type wing has been investigated in stalls of varying abruptness over the Mach number range from .18 to .49 and Reynolds number range from 6.1 to 13.4 million. The maximum lift coefficients were found to increase linearly with increasing rate of change of angle of attack per chord length of travel up to the maximum rate attained in the tests (.66 deg per chord length of travel) in contradistinction to the results of the flight tests of two other airplanes. The tests indicated that the Mach and Reynolds numbers effects were of sufficient importance to produce more than a twofold variation in the increment of Climax due to a given rate of change of angle of attack.