Engineering topics
Publications and source records attributed to Smith, Mark S..
This presentation increases awareness of the SAW project and the Convergent Aeronautics Solutions project by showing aerodynamic predictions and flight test results for the small-scale PTERA airplane.
This presentation describes aerodynamics and performance X-plane airworthiness guidelines and best practices.
Discuss ACTE aerodynamic modeling efforts and provide comparisons of predictions to flight results for lift and pitching moment increments.
A pair of compliant trailing edge flaps was flown on a modified GIII airplane. Prior to flight test, multiple analysis tools of various levels of complexity were used to predict the aerodynamic effects of the flaps. Vortex lattice, full potential flow, and full Navier-Stokes aerodynamic analysis software programs were used for prediction, in addition to another program that used empirical data. After the flight-test series, lift and pitching moment coefficient increments due to the flaps were estimated from flight data and compared to the results of the predictive tools. The predicted lift increments matched flight data well for all predictive tools for small flap deflections. All tools over-predicted lift increments for large flap deflections. The potential flow and Navier-Stokes programs predicted pitching moment coefficient increments better than the other tools.
The aerodynamic effects of compliant flaps installed onto a modified Gulfstream III airplane were investigated. Analyses were performed prior to flight to predict the aerodynamic effects of the flap installation. Flight tests were conducted to gather both structural and aerodynamic data. The airplane was instrumented to collect vehicle aerodynamic data and wing pressure data. A leading-edge stagnation detection system was also installed. The data from these flights were analyzed and compared with predictions. The predictive tools compared well with flight data for small flap deflections, but differences between predictions and flight estimates were greater at larger deflections. This paper describes the methods used to examine the aerodynamics data from the flight tests and provides a discussion of the flight-test results in the areas of vehicle aerodynamics, wing sectional pressure coefficient profiles, and air data.
No abstract available
As part of the Stratospheric Observatory for Infrared Astronomy program, a 747SP airplane was modified to carry a 2.5-m telescope in the aft section of the fuselage. The resulting airborne observatory allows for observations above 99 percent of the water vapor in the atmosphere. The open cavity created by the modifications had the potential to significantly affect the airplane in the areas of aerodynamics and acoustics. Several series of flight tests were conducted to clear the operating envelope of the airplane for astronomical observations, planned to be performed between the altitudes of 35,000 ft and 45,000 ft. The flight tests were successfully completed. Cavity acoustics were below design limits, and the overall acoustic characteristics of the cavity were better than expected. The modification did have some effects on the stability and control of the airplane, but these effects were not significant. Airplane air data systems were not affected by the modifications. This paper describes the methods used to examine the aerodynamics and acoustic data from the flight tests and provides a discussion of the flight-test results in the areas of cavity acoustics, stability and control, and air data.
As part of the Stratospheric Observatory for Infrared Astronomy (SOFIA) program, a 747SP airplane was modified to carry a 2.5 meter telescope in the aft section of the fuselage. The resulting airborne observatory allows for observations above 99 percent of the water vapor in the atmosphere. The open cavity created by the modifications had the potential to significantly affect the airplane in the areas of aerodynamics and acoustics. Several series of flight tests were conducted to clear the airplanes operating envelope for astronomical observations, planned to be performed between the altitudes of 39,000 feet and 45,000 feet. The flight tests were successfully completed. Cavity acoustics were below design limits, and the overall acoustic characteristics of the cavity were better than expected. The modification did have some effects on the stability and control of the airplane, but these effects were not significant. Airplane air data systems were not affected by the modifications. This paper describes the methods used to examine the aerodynamics and acoustic data from the flight tests and provides a discussion of the flight test results in the areas of cavity acoustics, stability and control, and air data.
The HIFiRE-1 flight experiment provided a valuable database pertaining to boundary layer transition over a 7-degree half-angle, circular cone model from supersonic to hypersonic Mach numbers, and a range of Reynolds numbers and angles of attack. This paper reports selected findings from the ongoing computational analysis of the measured in-flight transition behavior. Transition during the ascent phase at nearly zero degree angle of attack is dominated by second mode instabilities except in the vicinity of the cone meridian where a roughness element was placed midway along the length of the cone. The growth of first mode instabilities is found to be weak at all trajectory points analyzed from the ascent phase. For times less than approximately 18.5 seconds into the flight, the peak amplification ratio for second mode disturbances is sufficiently small because of the lower Mach numbers at earlier times, so that the transition behavior inferred from the measurements is attributed to an unknown physical mechanism, potentially related to step discontinuities in surface height near the locations of a change in the surface material. Based on the time histories of temperature and/or heat flux at transducer locations within the aft portion of the cone, the onset of transition correlated with a linear N-factor, based on parabolized stability equations, of approximately 13.5. Due to the large angles of attack during the re-entry phase, crossflow instability may play a significant role in transition. Computations also indicate the presence of pronounced crossflow separation over a significant portion of the trajectory segment that is relevant to transition analysis. The transition behavior during this re-entry segment of HIFiRE-1 flight shares some common features with the predicted transition front along the elliptic cone shaped HIFiRE-5 flight article, which was designed to provide hypersonic transition data for a fully 3D geometric configuration. To compare and contrast the crossflow dominated transition over the HIFiRE-1 and HIFiRE-5 configurations, this paper also analyzes boundary layer instabilities over a subscale model of the HIFiRE-5 flight configuration that was tested in the Mach 6 quiet tunnel facility at Purdue University.
The HIFiRE-1 flight experiment provided a valuable database pertaining to boundary layer transition over a 7-degree half-angle, circular cone model from supersonic to hypersonic Mach numbers, and a range of Reynolds numbers and angles of incidence. This paper reports the initial findings from the ongoing computational analysis pertaining to the measured in-flight transition behavior. Transition during the ascent phase at nearly zero degree angle of attack is dominated by second mode instabilities except in the vicinity of the cone meridian where a roughness element was placed midway along the length of the cone. The first mode instabilities were found to be weak at all trajectory points analyzed from the ascent phase. For times less than approximately 18.5 seconds into the flight, the peak amplification ratio for second mode disturbances is sufficiently small because of the lower Mach numbers at earlier times, so that the transition behavior inferred from the measurements is attributed to an unknown physical mechanism, potentially related to step discontinuities in surface height near the locations of a change in the surface material. Based on the time histories of temperature and/or heat flux at transducer locations within the aft portion of the cone, the onset of transition correlated with a linear PSE N-factor of approximately 14.
Practical aspects of identifying dynamic models for aircraft in real time were studied. Topics include formulation of an equation-error method in the frequency domain to estimate non-dimensional stability and control derivatives in real time, data information content for accurate modeling results, and data information management techniques such as data forgetting, incorporating prior information, and optimized excitation. Real-time dynamic modeling was applied to simulation data and flight test data from a modified F-15B fighter aircraft, and to operational flight data from a subscale jet transport aircraft. Estimated parameter standard errors, prediction cases, and comparisons with results from a batch output-error method in the time domain were used to demonstrate the accuracy of the identified real-time models.
Practical aspects of identifying dynamic models for aircraft in real time were studied. Topics include formulation of an equation-error method in the frequency domain to estimate non-dimensional stability and control derivatives in real time, data information content for accurate modeling results, and data information management techniques such as data forgetting, incorporating prior information, and optimized excitation. Real-time dynamic modeling was applied to simulation data and flight test data from a modified F-15B fighter aircraft, and to operational flight data from a subscale jet transport aircraft. Estimated parameter standard errors and comparisons with results from a batch output-error method in the time domain were used to demonstrate the accuracy of the identified real-time models.
An experimental multisegmented telescoping nose boom has been installed on an F-15B airplane to be tested in a flight environment. The experimental nose boom is representative of one that could be used to tailor the sonic boom signature of an airplane such as a supersonic business jet. The nose boom consists of multiple sections and could be extended during flight to a length of 24 ft. The preliminary analyses indicate that the addition of the experimental nose boom could adversely affect vehicle flight characteristics and air data systems. Before the boom was added, a series of flights was conducted to update the aerodynamic model and characterize the air data systems of the baseline airplane. The baseline results have been used in conjunction with estimates of the nose boom's influence to prepare for a series of research flights conducted with the nose boom installed. Data from these flights indicate that the presence of the experimental boom reduced the static pitch and yaw stability of the airplane. The boom also adversely affected the static-position error of the airplane but did not significantly affect angle-of-attack or angle-of-sideslip measurements. The research flight series has been successfully completed.
An experimental multisegmented telescoping nose boom has been installed on an F-15B airplane to be tested in a flight environment. The experimental nose boom is representative of one that could be used to tailor the sonic boom signature of an airplane such as a supersonic business jet. The nose boom consists of multiple sections and could be extended during flight to a length of 24 ft. The preliminary analyses indicated that the addition of the experimental nose boom could adversely affect vehicle flight characteristics and air data systems. Before the boom was added, a series of flights was flown to update the aerodynamic model and characterize the air data systems of the baseline airplane. The baseline results have been used in conjunction with estimates of the nose boom s influence to prepare for a series of research flights conducted with the nose boom installed. Data from these flights indicate that the presence of the experimental boom reduced the static pitch and yaw stability of the airplane. The boom also adversely affected the static-position error of the airplane but did not significantly affect angle-of-attack or angle-of-sideslip measurements. The research flight series has been successfully completed.
Photogrammetric analysis of high-speed digital video data was performed to estimate trajectories of foam debris ejected from an F-15B aircraft. This work was part of a flight test effort to study the transport properties of insulating foam shed by the Space Shuttle external tank during ascent. The conical frustum-shaped pieces of debris, called "divots," were ejected from a flight test fixture mounted underneath the F-15B aircraft. Two onboard cameras gathered digital video data at two thousand frames per second. Time histories of divot positions were determined from the videos post flight using standard photogrammetry techniques. Divot velocities were estimated by differentiating these positions with respect to time. Time histories of divot rotations were estimated using four points on the divot face. Estimated divot position, rotation, and Mach number for selected cases are presented. Uncertainty in the results is discussed.
Aerodynamic parameters were estimated based on flight data from the third flight of the X-43A hypersonic research vehicle, also called Hyper-X. Maneuvers were flown using multiple orthogonal phase-optimized sweep inputs applied as simultaneous control surface perturbations at Mach 8, 7, 6, 5, 4, and 3 during the vehicle descent. Aerodynamic parameters, consisting of non-dimensional longitudinal and lateral stability and control derivatives, were estimated from flight data at each Mach number. Multi-step inputs at nearly the same flight conditions were also flown to assess the prediction capability of the identified models. Prediction errors were found to be comparable in magnitude to the modeling errors, which indicates accurate modeling. Aerodynamic parameter estimates were plotted as a function of Mach number, and compared with estimates from the pre-flight aerodynamic database, which was based on wind-tunnel tests and computational fluid dynamics. Agreement between flight estimates and values computed from the aerodynamic database was excellent overall.
Near real-time stability and control derivative extraction is required to support flight demonstration of Intelligent Flight Control System (IFCS) concepts being developed by NASA, academia, and industry. Traditionally, flight maneuvers would be designed and flown to obtain stability and control derivative estimates using a postflight analysis technique. The goal of the IFCS concept is to be able to modify the control laws in real time for an aircraft that has been damaged in flight. In some IFCS implementations, real-time parameter identification (PID) of the stability and control derivatives of the damaged aircraft is necessary for successfully reconfiguring the control system. This report investigates the usefulness of Prescribed Simultaneous Independent Surface Excitations (PreSISE) to provide data for rapidly obtaining estimates of the stability and control derivatives. Flight test data were analyzed using both equation-error and output-error PID techniques. The equation-error PID technique is known as Fourier Transform Regression (FTR) and is a frequency-domain real-time implementation. Selected results were compared with a time-domain output-error technique. The real-time equation-error technique combined with the PreSISE maneuvers provided excellent derivative estimation in the longitudinal axis. However, the PreSISE maneuvers as presently defined were not adequate for accurate estimation of the lateral-directional derivatives.
An indirect adaptive system has been constructed for robust control of an aircraft with uncertain aerodynamic characteristics. This system consists of a multilayer perceptron pre-trained neural network, online stability and control derivative identification, a dynamic cell structure online learning neural network, and a model following control system based on the stochastic optimal feedforward and feedback technique. The pre-trained neural network and model following control system have been flight-tested, but the online parameter identification and online learning neural network are new additions used for in-flight adaptation of the control system model. A description of the modification and integration of these two stand-alone software packages into the complete system in preparation for initial flight tests is presented. Open-loop results using both simulation and flight data, as well as closed-loop performance of the complete system in a nonlinear, six-degree-of-freedom, flight validated simulation, are analyzed. Results show that this online learning system, in contrast to the nonlearning system, has the ability to adapt to changes in aerodynamic characteristics in a real-time, closed-loop, piloted simulation, resulting in improved flying qualities.