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The accuracy of the instruments used for flight test measurements has been inadequate so a new series of tests were commissioned with the development of an accurate elevator setting recorder.
Older commercial aircraft often exceed FAA (Federal Aviation Administration) sideline noise regulations. The major problem is the jet noise associated with the high exhaust velocities of the low bypass ratio engines on such aircraft. Mixer/ejector exhaust systems can provide a simple means of reducing the jet noise on these aircraft by mixing cool ambient air with the high velocity engine gases before they are exhausted to ambient. This paper presents new information on thrust performance predictions, and thrust augmentation capabilities of mixer/ejectors. Results are presented from the recent development program of the patented Alternating Lobe Mixer Ejector Concept (ALMEC) suppressor system for the Gulfstream GII, GIIB and GIII aircraft. Mixer/ejector performance procedures are presented which include classical control volume analyses, compound compressible flow theory, lobed nozzle loss correlations and state of the art computational fluid dynamic predictions. The mixer/ejector thrust predictions are compared to subscale wind tunnel test model data and actual aircraft flight test measurements. The results demonstrate that a properly designed mixer/ejector noise suppressor can increase effective engine bypass ratio and generate large thrust gains at takeoff conditions with little or no thrust loss at cruise conditions. The cruise performance obtained for such noise suppressor systems is shown to be a strong function of installation effects on the aircraft.
Time-dependent Reynolds-averaged Navier-Stokes simulations have been carried out for a NACA 0012 airfoil and a flexible UH-60A rotor undergoing dynamic stall. The OVERFLOW computational fluid dynamics (CFD) code is used to explore the use of near-body adaptive mesh refinement (NB-AMR) for the first time on a flexible helicopter rotor in forward flight. Emphasis is placed on understanding the temporal and spatial convergence of the solutions, their numerical efficiency and stability, and establishing a grid-converged solution. Some of the remaining differences between CFD and flight-test measurements are discussed. Time-dependent flow visualization is used to provide an improved understanding of the physical mechanisms involved with two-dimensional and three-dimensional dynamic stall, and the NB-AMR process.
Microphone phased-array and pole-mounted microphone data gathered during the NASA Acoustics Research Measurements flight tests were used to benchmark results from companion full-scale aeroacoustics simulations. Conducted with the lattice Boltzmann solver PowerFLOW®, the simulations predicted the acoustic behavior of various tested aircraft configurations. Emphasis was placed on those flown during the third flight test - a Fowler flap-equipped Gulfstream G-III with and without noise abatement technology on the main landing gear. Direct comparisons between experimental and synthetic microphone phasedarray data were achieved by applying the same processing and deconvolution technique to both sets of data. To extend the validation of the computations to the metric used for noise certification, the Effective Perceived Noise Level, a high-fidelity digital model of the nose landing gear, which was excluded from earlier computations, was developed and integrated into the G-III aircraft geometry. The acoustic study presented here demonstrates that the simulated beamform maps and corresponding integrated farfield spectra accurately predict the locations and strengths of the prominent airframe noise sources present on the G-III aircraft.
A preliminary field test of the accuracy of wind velocity measurements obtained using global positioning system-tracked rawinsonde balloons has been performed. Wind comparisons have been conducted using global positioning system (GPS) and radio automatic theodolite sounder (RATS) rawinsondes and a high-precision range instrumentation radar-tracked reflector. Wind velocity differences between the GPS rawinsondes and the radar were significantly less than between the RATS rawinsondes and the radar. These limited test results indicate a root-mean-square wind velocity difference from 4.98 kn (2.56 m/sec) for the radar and RATS to 1.09 kn (0.56 m/sec) for the radar and GPS. Differences are influenced by user reporting requirements, data processing techniques, and the inherent tracking accuracies of the system. This brief field test indicates that the GPS sounding system tracking data are more precise than the RATS system. When high-resolution wind data are needed, use of GPS rawinsonde systems can reduce the burden on range radar operations.
A preliminary field test of the accuracy of wind velocity measurements obtained using global positioning system-tracked rawinsonde balloons has been performed. Wind comparisons have been conducted using global positioning system (GPS) and radio automatic theodolite sounder (RATS) rawinsondes and a high-precision range instrumentation radar-tracked reflector. Wind velocity differences between the GPS rawinsondes and the radar were significantly less than between the RATS rawinsondes and the radar. These limited test results indicate a root-mean-square wind velocity difference from 4.98 kn (2.56 m/sec) for the radar and RATS to 1.09 kn (0.56 m/sec) for the radar and GPS. Differences are influenced by user reporting requirements, data processing techniques, and the inherent tracking accuracies of the system. This brief field test indicates that the GPS sounding system tracking data are more precise than the RATS system. When high-resolution wind data are needed, use of GPS rawinsonde systems can reduce the burden on range radar operations.
Flight test results for two sensor fault-tolerant algorithms developed for a redundant strapdown inertial measurement unit are presented. The inertial measurement unit (IMU) consists of four two-degrees-of-freedom gyros and accelerometers mounted on the faces of a semi-octahedron. Fault tolerance is provided by edge vector test and generalized likelihood test algorithms, each of which can provide dual fail-operational capability for the IMU. To detect the wide range of failure magnitudes in inertial sensors, which provide flight crucial information for flight control and navigation, failure detection and isolation are developed in terms of a multi level structure. Threshold compensation techniques, developed to enhance the sensitivity of the failure detection process to navigation level failures, are presented. Four flight tests were conducted in a commercial transport-type environment to compare and determine the performance of the failure detection and isolation methods. Dual flight processors enabled concurrent tests for the algorithms. Failure signals such as hard-over, null, or bias shift, were added to the sensor outputs as simple or multiple failures during the flights. Both algorithms provided timely detection and isolation of flight control level failures. The generalized likelihood test algorithm provided more timely detection of low-level sensor failures, but it produced one false isolation. Both algorithms demonstrated the capability to provide dual fail-operational performance for the skewed array of inertial sensors.
Flight test results of a vector-based fault-tolerant algorithm for a redundant strapdown inertial measurement unit are presented. Because the inertial sensors provide flight-critical information for flight control and navigation, failure detection and isolation is developed in terms of a multi-level structure. Threshold compensation techniques for gyros and accelerometers, developed to enhance the sensitivity of the failure detection process to low-level failures, are presented. Four flight tests, conducted in a commercial transport type environment, were used to determine the ability of the failure detection and isolation algorithm to detect failure signals, such a hard-over, null, or bias shifts. The algorithm provided timely detection and correct isolation of flight control- and low-level failures. The flight tests of the vector-based algorithm demonstrated its capability to provide false alarm free dual fail-operational performance for the skewed array of inertial sensors.
A theoretical and experimental investigation of methods for measuring the performance of general aviation airplanes was conducted using relatively simple instrumentation currently available and data extraction techniques established from efforts in other disciplines. The possibilities of improving flight test data by use of improved modern instrumentation and digital data recording and data analysis were considered.
The results of supersonic wind-tunnel tests on three probes at nominal Mach numbers of 1.6, 1.8 and 2.0 and flight tests on two of these probes up to a Mach number of 1.9 are described. One probe is an 8 deg. half-angle wedge with two total-pressure measurements and one static. The second, a conical probe, is a cylinder that has a 15 deg., semi-angle cone tip with one total-pressure orifice at the apex and four static-pressure orifices on the surface of the cone, 90 deg. apart, and about two-thirds of the distance from the cone apex to the base of the cone. The third is a 2 deg. semi-angle cone that has two static ports located 180 deg. apart about 1.5 inches behind the apex of the cone. The latter probe was included since it has been the "probe of choice" for wind-tunnel flow-field pressure measurements (or one similar to it) for the past half-century. The wedge and 15 deg. conical probes used in these tests were designed for flight diagnostic measurements for flight Mach numbers down to 1.35 and 1.15 respectively, and have improved capabilities over earlier probes of similar shape. The 15. conical probe also has a temperature sensor that is located inside the cylindrical part of the probe that is exposed to free-stream flow through an annulus at the apex of the cone. It enables the determination of free-stream temperature, density, speed of sound, and velocity, in addition to free-stream pressure, Mach number, angle of attack and angle of sideslip. With the time-varying velocity, acceleration can be calculated. Wind-tunnel tests of the two probes were made in NASA Langley Research Center fs Unitary Plan Wind Tunnel (UPWT) at Mach numbers of 1.6, 1.8, and 2.0. Flight tests were carried out at the NASA Dryden Flight Research Center (DFRC) on its F-15B aircraft up to Mach numbers of 1.9. The probes were attached to a fixture, referred to as the Centerline Instrumented Pylon (CLIP), under the fuselage of the aircraft. Problems controlling the velocity of the flow through the conical probe required for accurate temperature measurements are noted, as well as some calibration problems of the miniature pressure sensors that required a re-calculation of the flow variables. Data are presented for angle of attack, pressure and Mach number obtained in the wind tunnel and in flight. In the wind tunnel some transient data were obtained by translating the probes through the shock flow field created by a bump on the wind-tunnel wall.
Phased microphone array measurements obtained during flight tests conducted in 2016 and 2017 are used to assess the importance of local meteorological measurements on the data. In particular, the effectiveness of atmospheric absorption corrections is evaluated under vastly different temperature and humidity conditions. The results indicate that, even under conditions with high absorption, sources can be visualized up to a frequency that is dependent on background noise levels, wind, and atmospheric turbulence. However, absolute levels were found to be problematic on days with high absorption rates, with the discrepancies most prevalent for aircraft positions further from the center of the array. Restricting the data to those days with favorable meteorological conditions generally resulted in a good collapse of the spectra, with differences less than a couple of decibels.
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Postflight tracking system measurement error evaluation using regression analysis for Apollo Saturn 1B flight tests
Flight tests of radar scattering coefficient measuring instrument
The space transportation system 1 mission will be the first opportunity to assess the induced environment of the orbiter payload bay region. Two tools were developed to aid in this assessment. The shuttle payload contamination evaluation computer program was developed to provide an analytical tool for prediction of the induced molecular contamination environment of the space shuttle orbiter during its onorbit operations. An induced environment contamination monitor was constructed and tested to measure the space shuttle orbiter contamination environment inside the payload bay during ascent and descent and inside and outside the payload bay during the onorbit phase. Measurements are to be performed during the four orbital flight test series. Measurements planned for the first flight are described and predicted environmental data are discussed. The results indicate that the expected data are within the measurement range of the induced environment contamination monitor instruments evaluated, and therefore it is expected that useful contamination environmental data will be available after the first flight.
As the space industry continues to strive for more efficient launch vehicles, it must relyon increasingly accurate predictive models. Verification of models typically requires physical testing. Flight data measurements offer the most real and therefore the most accurate data for model correlation. As NASA prepares for the inaugural launch of Space Launch System(SLS), Artemis-1, they must rely heavily on predictive system models to ensure flight safety.NASA has implemented a Development Flight Instrumentation (DFI) system in hopes of recovering useful flight data to aid in model correlation. Historically, some of the most important flight measurements are those that monitor the potentially destructive dynamic interaction of the structural and propellant modes – a phenomenon known as pogo. Pogo is a dynamic instability that can occur on a launch vehicle during any phase of ascent. During this investigation an end-to-end assessment of the Artemis-I Pogo-related DFI was performed to identify any obstacles inherent in the current instrumentation system which may prevent successful measurement of the data necessary to validate the current predictive fluid-structural models. Input drive signals were approximated and applied to a state-space system model to derive predicted pressure and acceleration responses. These predicted responses were then fed through a simulation of the data acquisition process in order to recover predicted measurements. Finally, a mock Flight Data Analysis (FDA) was performed to assess the ability of these measurements to meet the Flight Test Objectives(FTO).