Aircraft Fault Detection Using Real-Time Frequency Response Estimation

In support of NASA's Fly-By-Light/Power-By-Wire (FBL/PBW) program, a series of flight tests were conducted by NASA Langley Research Center in February, 1995. The NASA Boeing 757 was flown past known RF transmitters to measure both external and internal radiated fields. The aircraft was instrumented with strategically located sensors for acquiring data on shielding effectiveness and internal coupling. The data are intended to support computational and statistical modeling codes used to predict internal field levels of an electromagnetic environment (EME) on aircraft. The software was an integral part of the flight tests, as well as the data reduction process. The software, which provided flight test instrument control, data acquisition, and a user interface, executes on a Hewlett Packard (HP) 300 series workstation and uses HP VEEtest development software and the C programming language. Software tools were developed for data processing and analysis, and to provide a database organized by frequency bands, test runs, and sensors. This paper describes the data acquisition system on board the aircraft and concentrates on the software portion. Hardware and software interfaces are illustrated and discussed. Particular attention is given to data acquisition and data format. The data reduction process is discussed in detail to provide insight into the characteristics, quality, and limitations of the data. An analysis of obstacles encountered during the data reduction process is presented.

The NASA Flying Controls Testbed (FLiC) is a relatively small and inexpensive unmanned aerial vehicle developed specifically to test highly experimental flight control approaches. The most recent version of the FLiC is configured with 16 independent aileron segments, supports the implementation of C-coded experimental controllers, and is capable of fully autonomous flight from takeoff roll to landing, including flight test maneuvers. The test vehicle is basically a modified Army target drone, AN/FQM-117B, developed as part of a collaboration between the Aviation Applied Technology Directorate (AATD) at Fort Eustis, Virginia and NASA Langley Research Center. Several vehicles have been constructed and collectively have flown over 600 successful test flights, including a fully autonomous demonstration at the Association of Unmanned Vehicle Systems International (AUVSI) UAV Demo 2005. Simulations based on wind tunnel data are being used to further develop advanced controllers for implementation and flight test.

An airborne differential absorption lidar (DIAL) system has been developed at the NASA Langley Research Center for the remote measurement of water vapor (H2O) and aerosols in the lower troposphere. Significant modifications to the laser transmitters and other major subsystems have been implemented during the past two years to improve the system's performance and field reliability. The previous configuration of the airborne H2O DIAL system and the associated atmospheric data collected during flight tests at Wallops Flight Facility are described in detail in two recently published papers (Higdon et al., 1994 and Ponsardin et al., 1994). The modified system is to be flight tested in late 1994, and the system performance characteristics and preliminary atmospheric H2O and aerosol data from these flights are discussed in this paper.

Small unmanned aircraft systems (sUAS) face unique challenges in system identification due to their lightweight, small size, and increased sensitivity to wind and gusts. To address these challenges, flight test procedures for sUAS system identification using open-loop and closed-loop flight data are presented. The approaches are demonstrated through system identification of a small flying-wing UAS equipped with a low-cost Pixhawk autopilot. The open-loop method is demonstrated through identification of the longitudinal bare-airframe aircraft dynamics. Longitudinal frequency responses were estimated from the open-loop flight data, and a linear state space model was identified from the estimated frequency responses and supplemental trim data, which was used to better identify parameters related to the low-frequency phugoid mode. The identified phugoid mode matches well with low-frequency oscillations measured from flight data in the time domain. The closed-loop method is demonstrated through identification of the lateral-directional dynamics. This approach is shown to improve the signal-to-noise ratio and reduce deviation from the reference flight condition, which improves modeling results, and it can be used to simultaneously identify the bare-airframe, closed-loop, and broken-loop UAS dynamics from the same set of closed-loop flight data, reducing time and efforts spent flight testing. The identified lateral-directional model is compared with closed- and broken-loop frequency responses estimated from flight data. The results show excellent agreement, with all computed metrics, such as stability margins, matching within 10%, demonstrating the effectiveness of this method.

NASA Langley Research Center (LaRC) and The MITRE Corporation (MITRE) have developed - and successfully demonstrated - an integrated simulation-to-flight capability for evaluating sense and avoid (SAA) system elements. This integrated capability consists of a MITRE developed fast-time computer simulation for evaluating SAA algorithms, and a NASA LaRC surrogate unmanned aircraft system (UAS) equipped to support hardware and software in-the-loop evaluation of SAA system elements (e.g., algorithms, sensors, architecture, communications, autonomous systems), concepts, and procedures. The fast-time computer simulation subjects algorithms to simulated flight encounters/ conditions and generates a fitness report that records strengths, weaknesses, and overall performance. Reviewed algorithms (and their fitness report) are then transferred to NASA LaRC where additional (joint) airworthiness evaluations are performed on the candidate SAA system-element configurations, concepts, and/or procedures of interest; software and hardware components are integrated into the Surrogate UAS' research systems; and flight safety and mission planning activities are completed. Onboard the Surrogate UAS, candidate SAA system element configurations, concepts, and/or procedures are subjected to flight evaluations and in-flight performance is monitored. The Surrogate UAS, which can be controlled remotely via generic Ground Station uplink or automatically via onboard systems, operates with a NASA Safety Pilot/Pilot in Command onboard to permit safe operations in mixed airspace with manned aircraft. An end-to-end demonstration of a typical application of the capability was performed in non-exclusionary airspace in October 2011; additional research, development, flight testing, and evaluation efforts using this integrated capability are planned throughout fiscal year 2012 and 2013.

Liquid propellant makes up a significant portion of the total weight for large launch vehicles such as Saturn V, Space Shuttle, and the Space Launch System. Careful attention must be given to the influence of fuel slosh motion on the stability of the vehicle in the design of the Flight Control System (FCS). Historically, there have been instances where a poorly designed FCS in addition to a lack of passive damping have caused the slosh mass to drive the launch vehicle unstable. The dynamics behind controlling a quadcopter/hanging mass configuration is analogous to that of controlling the attitude of a rocket with a single propellant tank. The quadcopter/hanging mass configuration offers a reasonably accurate platform for assessing the real-time effectiveness of the SLS Adaptive Augmenting Controller in suppressing slosh instability. Flight test experiments were carried out at the NASA Langley Research Center's Autonomy Incubator. During both simulation and flight test, the hanging mass was intentionally made unstable and the adaptive algorithm successfully suppressed the instability as expected.

NASA has a technology program in place to enable the development of a next generation Reusable Launch Vehicle that will carry our future payloads into orbit at a much-reduced cost. The VentureStar, Lifting Body (LB) flight vehicle, is one of the potential reusable launch vehicle configurations being studied. A LB vehicle has no wings and derives its lift solely from the shape of its body, and has the unique advantages of superior volumetric efficiency, better aerodynamic efficiency at high angles-of-attack and hypersonic speeds, and reduced thermal protection system weight. Classically, in a ballistic vehicle, drag has been employed to control the level of deceleration in reentry. In the LB, lift enables the vehicle to decelerate at higher altitudes for the same velocity and defines the reentry corridor which includes a greater cross range. This paper outlines the flight stability and control aspects of our LB heritage which was utilized in the design of the VentureStar LB and its test version, the X-33. NASA and the U.S. Air Force have a rich heritage of LB vehicle design and flight experience. In the initial LB Program, eight LB's were built and over 225 LB test flights were conducted through 1975. Three LB series were most significant in the advancement of today's LB technolocy: the M2-F; the HL-10; and the X-24 series. The M2-F series was designed by NASA Ames Research Center, the HL-10 series by NASA Langley Research Center, and the X-24 series by the U. S. Air Force. LB vehicles are alive again today with the X- 33, X-38, and VentureStar.

NASA-Langley Research Center (NASA-LaRC) and MITRE have developed and successfully demonstrated an integrated simulation-to-flight capability for evaluating sense-and-avoid (SAA).The NASA-MITRE team has developed an integrated simulation and flight test capability for testing prototype sense-and-avoid system elements with an initial emphasis on ADS-B based algorithmic alternatives for self separation. This integrated system includes the capability to develop algorithms, perform fast-time simulation testing, generate fitness reports, integrate into the flight software framework, and execute flight tests in real-world conditions. The value of this integrated process was fully demonstrated and will be expanded to accommodate the flight evaluations scheduled for mid-2013 in eastern North Dakota. This process can serve as a model for the development of systems and algorithms needed to enable the safe integration of UAS in the NAS.

<div class="htmlview paragraph">A cooperative effort has been defined between NASA-Ames/Dryden Flight Research Facility and NASA-Langley Research Center in support of the X-29A Advanced Technology Demonstrator. The effort involves three phases: flight testing, wind-tunnel testing in the National Transonic Facility, and computational support of each experimental phase. These efforts are primarily aimed at understanding the complex flow phenomena and component interactions associated with the X-29A. Each phase of the effort is discussed in detail and initial data comparisons are presented. In summary, the synergistic effects of the complementary phases are identified, which will enhance the understanding of the unique aerodynamics of the X-29A.</div>

Aparallelcodeisdevelopedtosimulatenumericallywakevortexdetectionusingaradioacousticsoundingsystem (RASS). The code is written in FORTRAN 90 with the message passing interface for parallel implementation. The numerical simulation solves simultaneously the linearized Euler equations for a nonuniform mean e ow and the Maxwellequationsforanonhomogeneousmedium.Theradartransmitterandreceiverantennasaremodeledusing an array of point sources and a beam-forming technique, respectively. Many features of the RASS are explored using thenumericalsimulation.First,a uniformmeane owisconsidered,and theRASSsimulation isperformedfor two different types of incident acoustic e eld: a short single-frequency acoustic pulse and a continuous broadband acoustic source. Both monostatic and bistatic cone gurations are examined, and their results are compared. Taylor and Oseen vortex velocity proe les are used as samplemodels, and their mean e owe eldsare reconstructed from the backscattered electromagnetic signal using the Abel transform. The effect of radar beam width is also considered, as are the issues of nonaxisymmetric and interacting vortices. I. Introduction T HEcapacity of airports is constrainedseverelyby the airtrafe c control system’ s consideration of aircraft wake vortices. The present Instrument Flight Rules (IFR) restrictions are based on aircraft weight. The “ 3‐ 4‐ 5‐ 6 rule” sets the separation distances of aircraft by categories from small, less than 41,000 lb (18,600 kg), to heavy, greater than 255,000 lb (116,000 kg). These separations are viewed as very conservative. However, because there remains signie cant uncertainty about wake vortex behavior under different atmospheric conditions, there are considerable technological barriers to improvements in terminal area productivity. In an attempt to enhance airport capacity, an aircraft vortex spacing system (AVOSS) has been under development at NASA Langley Research Center. The AVOSS will provide the means to allow air trafe c control to reduce spacing safely in instrument operations when the appropriate weather conditions exist. A key element in AVOSS is the development of a real-time reliable wake vortex detection system (WVDS). A radio acoustic sounding system (RASS) is a promising candidate for a WVDS. The basic concept of RASS is based on the tracking of sound waves with radar. When acoustic waves are transmitted, they produce pressure perturbations, which lead to e uctuations in the atmospheric permittivity. An incident electromagnetic e eldisthenscatteredduetothepermittivityvariationsandgenerates an echo that can be detected by a receiver antenna. The spectrum of this echo shows a Doppler shift proportional to the local speed of sound. However, the backscattered signals are so weak that no measurable signal would be received unless the Bragg condition is satise ed.The Bragg condition is a relationshipbetween the incident electromagnetic and acoustic wavelengths that ensures a constructive interference between the two waves. The Bragg condition may be written as ¸e D 2¸a sin®

The Landing and Impact Research Facility (LandIR) at NASA Langley Research Center is a 240 ft. high A-frame structure which is used for full-scale crash testing of aircraft and rotorcraft vehicles. Because the LandIR provides a unique capability to introduce impact velocities in the forward and vertical directions, it is also serving as the facility for landing tests on full-scale and sub-scale Orion spacecraft mass simulators. Recently, a threedimensional photogrammetry system was acquired to assist with the gathering of vehicle flight data before, throughout and after the impact. This data provides the basis for the post-test analysis and data reduction. Experimental setups for pendulum swing tests on vehicles having both forward and vertical velocities can extend to 50 × 50 × 50 foot cubes, while weather, vehicle geometry, and other constraints make each experimental setup unique to each test. This paper will discuss the specific calibration techniques for large fields of views, camera and lens selection, data processing, as well as best practice techniques learned from using the large field of view photogrammetry on a multitude of crash and landing test scenarios unique to the LandIR.

As part of the Adaptive Vehicle Management Systems (AVMS) Phase II program, an Active External Load Stabilization (AELS) system was designed that consists of Flight Control Load Stabilization (FCLS) and Active Cargo Hook (ACH) methods to stabilize a load which can also be combined and operated simultaneously. Prior flight testing revealed issues with the performance of the AELS system that were a result of a modeling error as well as the need to utilize different underlying flight control system control law feedback gains when flying with an external load. Fixing the modeling error had a significant impact on the evaluation of the system in piloted simulation, particularly for the FCLS method. A more comprehensive control law gain optimization was employed using CONDUIT® in order to achieve all of the desired design requirements, such as stability margins, handling qualities (HQ), and load-damping (LD). Changes were also made to the mode switching between the FCLS-HQ and FCLS-LD modes. Following these updates, the system was re-evaluated in piloted simulations and flight test. The results show that the ACH and FCLS systems each performed extremely well; significantly reducing load motion and allowing the pilot to more quickly complete tasks with an external load. The combined ACH+FCLS system performed very well, however even better performance is expected with further refinements.

The effect of miniflaps for increasing the L/D ratio and the lift coefficient has been studied on airliners as well as on UAV-s and wind turbines. For sailplanes the lift when Cl > 1.0 is of main interest. As the maximum wing loading of racing sailplanes reaches 60–62 kg/m2, it is necessary to achieve a high Cl max (1.7–1.8) in thermals. In this case the decrease in TAS caused by a high Cl max even compensates for the drop of the L/D ratio to a certain extent, as the climb speed will increase when the spiral flight radius diminishes in thermals. To bring the L/D to Cl > 1.0, a 2% chord miniflap at a 30° deflection angle was attached to the trailing edge of a Jantar-Standard 3 type sailplane wing (airfoil NN-8). In flight tests it was found that the miniflap increased the sailplane‘s Cl max to 1.35–1.66, i.e. by 23% (Re 1.0–0.92×106). At the same time the L/D ratio Cl increased by over 1.0. Especially good L/D improvement was noted with Cl at 1.13–1.19. In thermal Cl of 1.57–1.65 the roll control was good. At lower Cl < 1.0 values, the miniflap reduced the L/D ratio in comparison with a normal configuration.

The 30x60 feet wind tunnel at the NASA Langley Research Center is a major facility capable of testing full size aircraft. Air pressure measurements recorded during aerodynamic tests of aircraft or models are made using a probe attached to a three axis carriage. Positioning of the carriage during tests is made using manual observations of survey tapes and an electronic counter. Observations with three theodolites were used to calibrate the location of the carriage relative to an orthogonal coordinate system with a vertical axis. Calibrations were carried out with both the air flow off and the air flow on to assess the influence of induced vibrations on the measurements and the displacement of the carriage under wind load. A variety of coordinate transformations were used to analyse the data, demonstrate the subjectivity of the manual observations and provide a mathematical correction for the observed data.

Flight test activities that expand the flight envelope are inherently expensive and at times dangerous, particularly when approaching stability and/or flutter boundaries. With the computational power now available, on-line software applications have emerged that can clear test points while a flight is in progress. For flying qualities evaluations, it is important to identify accurate frequency responses from which the parameters associated with commonly used criteria are obtained. Typically, these frequency responses are generated from long duration test inputs such as frequency sweeps that may be minutes long depending on the frequency range of interest. To meet goals set by the Air Force Flight Test Center, a new system identification method – the Narrowband Signature (NBS) – has been developed to obtain accurate frequency responses from short duration (five seconds or less) test inputs. A flying qualities flight test database was created using the Calspan Corporation Learjet In-Flight Simulator to exercise the new identification method. Comparisons were made against industry standard fast Fourier transform (FFT)-based identification techniques for both long and short duration inputs with a longitudinal axis cruise configuration that featured well-damped short period dynamics as well as a lateraldirectional approach configuration that featured a lightly damped Dutch roll mode. The identification results for the cruise configuration found that all methods identified the system accurately with minimal adjustment of the default identification parameters for the long duration input. For those configurations with lightly damped modes, even with the long duration inputs, adjustment of the recommended window sizes was necessary for proper identification of the lightly damped mode. For the short duration input cases, NBS showed a clear advantage over the FFT-based techniques for both configurations. To demonstrate accuracy of the methods, the results are compared against transfer function models of the selected configurations.