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The quasi-satellite (QS) phenomenon makes two celestial bodies to fly near each other (Mikkola et al. 2006) and that effect can be used also to make artificial satellites move in tandem. We consider formation flight of two or three satellites in low eccentricity near Earth orbits. With the help of weak ion thrusters it is possible to accomplish tandem flight. With ion thrusters it is also possible to mimic many kinds of mutual force laws between the satellites. We found that both a constant repulsive force or an attractive force that decreases with the distance are able to preserve the formation in which the eccentricities cause the actual relative motion and the weak thrusters keep the mean longitude difference small. Initial values are important for the formation flight but very exact adjustment of orbital elements is not important. Simplicity is one of our goals in this study and this result is achieved at least in the way that, when constant force thrusters are used, the satellites only need to detect the directions of the other ones to fly in tandem. A repulsive acceleration of the order of 10^-6 times the Earth attraction, is enough to effectively eliminate the disruptive effects of all the perturbations at least for a timescale of years.

5% scale Blended-Wing-Body free-flight model in the Langley Full Scale Tunnel, building 640. People in the photographs include Sue Grafton (with model), Dan Vicroy (with model). Group photo: S. Grafton, D. Vicroy, D. Murri, C. Buttrill, J. Brandon, M. Croom, B. Jackson, C. Debro, G. Adams.

5% scale Blended-Wing-Body free-flight model in the Langley Full Scale Tunnel, building 640. People in the photographs include Sue Grafton (with model), Dan Vicroy (with model). Group photo: S. Grafton, D. Vicroy, D. Murri, C. Buttrill, J. Brandon, M. Croom, B. Jackson, C. Debro, G. Adams.

5% scale Blended-Wing-Body free-flight model in the Langley Full Scale Tunnel, building 640. People in the photographs include Sue Grafton (with model), Dan Vicroy (with model). Group photo: S. Grafton, D. Vicroy, D. Murri, C. Buttrill, J. Brandon, M. Croom, B. Jackson, C. Debro, G. Adams.

Photographed on 05/17/2005. -- Group photographs of: 1) Flight Dyanmics Branch (FDB), 2) Dynamic Systems and Controls Branch (DSCB) and 3) Combined FDB and DSCB groups.

Thesis (M.A. in A.E.)--Naval Postgraduate School, 1967

5% scale Blended-Wing-Body free-flight model in the Langley Full Scale Tunnel, building 640. People in the photographs include Sue Grafton (with model), Dan Vicroy (with model). Group photo: S. Grafton, D. Vicroy, D. Murri, C. Buttrill, J. Brandon, M. Croom, B. Jackson, C. Debro, G. Adams.

The application of GPS to spacecraft attitude determination is a new and growing field. Although the theoretical literature is extensive, space flight testing is currently sparse and inadequate. As an operations organization, the Flight Dynamics Division (FDD) has the responsibility to investigate this new technology, and determine how best to implement the innovation to provide adequate support for future missions. This paper presents some of the current efforts within FDD with regard to GPS attitude determination. This effort specifically addresses institutional capabilities to accommodate a new type of sensor, critically evaluating the literature for recent advancements, and in examining some available -albeit crude- flight data.

An aeroelastic analysis of the behavior of an entry vehicle utilizing an attached inflatable aerodynamic decelerator during supersonic flight is presented. The analysis consists of a planar, four degree of freedom simulation. The aeroshell and the IAD are assumed to be separate, rigid bodies connected with a spring-damper at an interface point constraining the relative motion of the two bodies. Aerodynamic forces and moments are modeled using modified Newtonian aerodynamics. The analysis includes the contribution of static aerodynamic forces and moments as well as pitch damping. Two cases are considered in the analysis: constant velocity flight and planar free flight. For the constant velocity and free flight cases with neutral pitch damping, configurations with highly-stiff interfaces exhibit statically stable but dynamically unstable aeroshell angle of attack. Moderately stiff interfaces exhibit static and dynamic stability of aeroshell angle of attack due to damping induced by the pitch angle rate lag between the aeroshell and IAD. For the free-flight case, low values of both the interface stiffness and damping cause divergence of the aeroshell angle of attack due to the offset of the IAD drag force with respect to the aeroshell center of mass. The presence of dynamic aerodynamic moments was found to influence the stability characteristics of the vehicle. The effect of gravity on the aeroshell angle of attack stability characteristics was determined to be negligible for the cases investigated.

View from the groundside end of the tether of the Makani M600 in flight. The camera captures tether dynamics during the second unconstrained flight of the M600—Remote Perch Crosswind Flight 02 (RPX-02), December 14, 2016. For more information, please see "Selected Decks from RPX Lessons Learned Reviews," included in The Energy Kite: Selected Results From the Design, Development and Testing of Makani’s Airborne Wind Turbines, Part III, and available at x.company/projects/makani.

Contents: Concepts for aeroelastic system approximations; static aerodynamics for flutter analyses; flutter at high mach numbers; an indicial flutter analysis for hypersonic delta wings; a theory for aeroelastic studies of delta lifting surfaces; flutter of flat panels in a low supersonic flow; flutter of rectangular panels; model flight testing on high-speed tracks; liquid behavior in rocket propellant tanks; dynamics of LP vehicles; aero-inertial control system; dynamic loads of missile configurations; wind loads on a vertically rising vehicle; random gust and taxi response calculations for delta wing aircraft; blast-loading on airfoils; stall buffeting loads; a method for analyzing heated wings; deformational response of heated wing structures; thermal stiffness; acoustic fatigue tests for elevated temperatures structural design; structural vibration in space vehicles; structural response in noise inputs; captive missile response to random pressures; structural response to the noise input of the Saturn engines; the environmental vibration problem.

The 1987 NASA/Army Rotorcraft Technology Conference was held at Ames Research Center, Moffett Field, California, March 17-19, 1987. The Conference Proceedings is a compilation of over 30 technical papers presented at this milestone event which reported on the advances in rotorcraft technical knowledge resulting from NASA, Army, and industry rotorcraft research programs over the last 5 to 10 years. The Conference brought together over 230 government, industry, and allied nation conferees to exchange technical information and hear invited technical papers by prominent NASA, Army, and industry researchers covering technology topics including aerodynamics, dynamics and aeroelasticity, propulsion and drive systems, flight dynamics and control, acoustics, systems integration, and research aircraft.

This is the Flight Dynamics portion of the generic Modular Simulator System (MSS) specification. It is designed to be tailored to specify the requirements for a specific aircraft training device or family of aircraft training devices. This specification contains specific tailoring instructions for each paragraph. When the tailoring process complete, the italicized tailoring instructions should have been replaced by application specific text or deleted from the specification. It is suggested that the user read the 'Modular simulator Engineering Guide' and the 'Modular Simulator Management Guide' prior to tailoring this volume. Modular Simulator System (MSS).

Moments of inertia were experimentally determined and longitudinal and lateral/directional static and dynamic stability and control derivatives were estimated for a fixed wing Unmanned Air Vehicle (UAV). Dynamic responses to various inputs were predicted based upon the estimated derivatives. A divergent spiral mode was revealed, but no particularly hazardous dynamics were predicted. The aircraft was then instrumented with an airspeed indicator, which when combined with the ability to determine elevator deflection through trim setting on the flight control transmitter, allowed for the determination of the aircraft's neutral point through flight test. The neutral point determined experimentally corresponded well to the theoretical neutral point. However, further flight testing with improved instrumentation is planned to raise the confidence level in the neutral point location. Further flight testing will also include dynamic studies in order to refine the estimated stability and control derivatives. Unmanned Aerial Vehicle, UAV, Flight test, Flight dynamics

Survey of numerical solution techniques for variational problems in rocket flight dynamics - U.S.S.R.

CHARACTERISTICS OF AIR MOTION WITH RESPECT TO GROUND AND ANALYSIS OF DYNAMICS OF AIRCRAFT IN TURBULENT AIR

The state of hingless rotorcraft research and development in the NATO countries as of 1973 is described. The scope of this report is limited to flight dynamics (as defined in the Preface) since most of the hingeless rotorcraft problems have occurred in this area. In the Introduction, the special place of the hingeless rotorcraft within the family of rotorcraft is considered. The chapter on the history of hingeless rotorcraft describes the hingeless rotor research and development of the various rotorcraft manufacturers and the hingless rotor research at government laboratories and universities. A hierarchy of dynamic concepts from isolated blade dynamics to complete rotor/ body dynamics is introduced. The effects of the basic rotor design parameters on flight dynamics are traced and certain hingeless rotorcraft problems are treated in some detail. A special chapter is devoted to the alleviation of hingeless rotor flight-dynamics problems by feedback control systems. Finaly, analytical modeling techniques, mathematical analysis techniques, and model and flight testing techniques for hingeless rotorcraft are discussed.

The purpose of this research was to evaluate the effectiveness of a team building organization development (OD) intervention currently underway at the Air Force Flight Dynamics Laboratory (AFFDL). Effectiveness was determined by analyzing the longitudinal change from 1978 to 1979 on eleven criterion variables: employee job satisfaction; job motivation; absenteeism; five factor dimensions of organizational climate; and for Scientists and Engineers (S and Es), three productivity factors. All variables except absenteeism were measured with a questionnaire. Absenteeism data was obtained from laboratory records. OD was found to have had some positive effect on AFFDL, but only for two (at most) of the eleven criteria. However, there was some evidence that the process of OD-induced change within the laboratory was still occurring despite the fact that the program was over three years old at the time of the second measurement. The apparent languid pace associated with organizational change observed in this study was attributed to two causes: the lack of an internal OD facilitator to lend some continuity to the program and the low-key emphasis and support for the program.

No Description

Report summarizes the findings of a five-year program devoted to improving fundamental knowledge on unsteady aerodynamic phenomena related to flight vehicles and on associated aeroelastic problems. With regard to minimum- weight structural optimization with aeroelastic constraints, both new results and new methods of solution for free and forced motion were published. The effect of chordwise-force components on flutter of large aspect ratio wings proved often to be unfavorable. Improved steady and unsteady theories were published for the loading of vertical-axis wind turbines, and discoveries were made regarding free vibration of their curved blades. It was learned how to adapt linear theory for simple harmonic oscillation to cover arbitrary small motion, with applications to automatic control. A nonlinear approach was published for transient lifting airloads at low speeds. A study was undertaken on aerodynamics useful for the analysis of variable-geometry propulsive devices. An approximate scheme was devised for highlighting the importance of partial- chord shocks for transonic aeroelastic stability, their influence proving often large and unfavorable.

Book digitized by Google from the library of the University of Michigan and uploaded to the Internet Archive by user tpb.

288 p. 24 cm

This thesis uses Flightlab to model and analyze the flight dynamics of the SH-6OB Seahawk helicopter. Flightlab runs on computers utilizing the UNIX operating system and is used for design, analysis and testing of an aircraft using non-linear modeling techniques. It is capable of modeling conventional main rotor-tail rotor and tandem rotor helicopters and tilt rotor aircraft. A procedural guide for modeling and analyzing a single main rotor helicopter is presented using the SH-60B. The non-linear response from lateral and longitudinal cyclic, main rotor collective and tail rotor collective inputs are presented. Flightlab is also capable of reducing the non-linear model to a linear model. The linear and non-linear models are then compared. The purpose of this thesis is to present a guide for using Flightlab to model and analyze an existing helicopter design, and also to have in place a well tested model to be used for further research.

5% scale Blended-Wing-Body free-flight model in the Langley Full Scale Tunnel, building 640. People in the photographs include Sue Grafton (with model), Dan Vicroy (with model). Group photo: S. Grafton, D. Vicroy, D. Murri, C. Buttrill, J. Brandon, M. Croom, B. Jackson, C. Debro, G. Adams.

5% scale Blended-Wing-Body free-flight model in the Langley Full Scale Tunnel, building 640. People in the photographs include Sue Grafton (with model), Dan Vicroy (with model). Group photo: S. Grafton, D. Vicroy, D. Murri, C. Buttrill, J. Brandon, M. Croom, B. Jackson, C. Debro, G. Adams.

We propose a new method to estimate the 6-dof trajectory of a flying object such as a quadrotor UAV within a 3D airspace monitored using multiple fixed ground cameras. It is based on a new structure from motion formulation for the 3D reconstruction of a single moving point with known motion dynamics. Our main contribution is a new bundle adjustment procedure which in addition to optimizing the camera poses, regularizes the point trajectory using a prior based on motion dynamics (or specifically flight dynamics). Furthermore, we can infer the underlying control input sent to the UAV's autopilot that determined its flight trajectory. Our method requires neither perfect single-view tracking nor appearance matching across views. For robustness, we allow the tracker to generate multiple detections per frame in each video. The true detections and the data association across videos is estimated using robust multi-view triangulation and subsequently refined during our bundle adjustment procedure. Quantitative evaluation on simulated data and experiments on real videos from indoor and outdoor scenes demonstrates the effectiveness of our method.

We propose a new method to estimate the 6-dof trajectory of a flying object such as a quadrotor UAV within a 3D airspace monitored using multiple fixed ground cameras. It is based on a new structure from motion formulation for the 3D reconstruction of a single moving point with known motion dynamics. Our main contribution is a new bundle adjustment procedure which in addition to optimizing the camera poses, regularizes the point trajectory using a prior based on motion dynamics (or specifically flight dynamics). Furthermore, we can infer the underlying control input sent to the UAV's autopilot that determined its flight trajectory. Our method requires neither perfect single-view tracking nor appearance matching across views. For robustness, we allow the tracker to generate multiple detections per frame in each video. The true detections and the data association across videos is estimated using robust multi-view triangulation and subsequently refined during our bundle adjustment procedure. Quantitative evaluation on simulated data and experiments on real videos from indoor and outdoor scenes demonstrates the effectiveness of our method.

We propose a new method to estimate the 6-dof trajectory of a flying object such as a quadrotor UAV within a 3D airspace monitored using multiple fixed ground cameras. It is based on a new structure from motion formulation for the 3D reconstruction of a single moving point with known motion dynamics. Our main contribution is a new bundle adjustment procedure which in addition to optimizing the camera poses, regularizes the point trajectory using a prior based on motion dynamics (or specifically flight dynamics). Furthermore, we can infer the underlying control input sent to the UAV's autopilot that determined its flight trajectory. Our method requires neither perfect single-view tracking nor appearance matching across views. For robustness, we allow the tracker to generate multiple detections per frame in each video. The true detections and the data association across videos is estimated using robust multi-view triangulation and subsequently refined during our bundle adjustment procedure. Quantitative evaluation on simulated data and experiments on real videos from indoor and outdoor scenes demonstrates the effectiveness of our method.

We propose a new method to estimate the 6-dof trajectory of a flying object such as a quadrotor UAV within a 3D airspace monitored using multiple fixed ground cameras. It is based on a new structure from motion formulation for the 3D reconstruction of a single moving point with known motion dynamics. Our main contribution is a new bundle adjustment procedure which in addition to optimizing the camera poses, regularizes the point trajectory using a prior based on motion dynamics (or specifically flight dynamics). Furthermore, we can infer the underlying control input sent to the UAV's autopilot that determined its flight trajectory. Our method requires neither perfect single-view tracking nor appearance matching across views. For robustness, we allow the tracker to generate multiple detections per frame in each video. The true detections and the data association across videos is estimated using robust multi-view triangulation and subsequently refined during our bundle adjustment procedure. Quantitative evaluation on simulated data and experiments on real videos from indoor and outdoor scenes demonstrates the effectiveness of our method.

This paper is in the public domain in USA. Metadata comes from the CrossRef API, see full record in the source URL below.

This paper describes model structures and parameter estimation algorithms suitable for the identification of unsteady aerodynamic models from input-output data. The model structures presented are state space models and include linear time-invariant (LTI) models and linear parameter-varying (LPV) models. They cover a wide range of local and parameter dependent identification problems arising in unsteady aerodynamics and nonlinear flight dynamics. We present a residue algorithm for estimating model parameters from data. The algorithm can incorporate apriori information and is described in detail. The algorithms are evaluated on the F-16XL wind-tunnel test data from NAS Langley Research Center. Results of numerical evaluation are presented. The paper concludes with a discussion major issues and directions for future work.

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The concept of rocket thrust stand simulation of space vehicle flight dynamics is developed. An electro-mechanical system is described wherein the space vehicle rocket motor attach point motion may be simulated. The physical dimensions of the system are small such that it may be used in conjunction with current environmental test cells. The degree of space vehicle simulation is a consequence of the limitations on the excursions of the active strut thrust stand. The resulting motion is shown to be vibratory with respect to the center of mass of the space vehicle system. Main emphasis is on the development of the analog simulator differential equations which ultimately serve as inputs to the mechanical system. These equations produce the vehicle motion as a function of the physically measurable thrust stand quantities, indicated thrust, and motor acceleration. The development of these equations consists of elimination of the steady-state component of vehicle motion and obtaining analog stability without interfering with the motor simulation. The resulting equations are presented in a general form so that simulation may be accomplished for any space vehicle whose dynamic modes may be described by time dependent, analog-solvable differential equations.

Moments of inertia were experimentally determined and longitudinal and lateral/directional static and dynamic stability and control derivatives were estimated for a fixed wing Unmanned Air Vehicle (UAV). Dynamic responses to various inputs were predicted based upon the estimated derivatives. A divergent spiral mode was revealed, but no particularly hazardous dynamics were predicted. The aircraft was then instrumented with an airspeed indicator, which when combined with the ability to determine elevator deflection through trim setting on the flight control transmitter, allowed for the determination of the aircraft's neutral point through flight test. The neutral point determined experimentally corresponded well to the theoretical neutral point. However, further flight testing with improved instrumentation is planned to raise the confidence level in the neutral point location. Further flight testing will also include dynamic studies in order to refine the estimated stability and control derivatives

While the linear theory has been the backbone of ballistic missile design and has served with excellent success in its various forms over the years, recent spectacular missile flight failures have appeared unexpectedly and are unaccounted for by this linear theory. It is these flight failures which now require attention and consideration of the non-linearities in the fluid force and moment system which contribute to missile flight performance and dynamic stability. Two types of non-linear flight instability have been isolated and identified as NonLinear Magnus Instability, and Catastrophic Yaw. It is these 2 flight instabilities which will be discussed. Approximate mathematical models will be suggested for the evaluation of missile dynamic stability. Both experimental and analytical applications of these models will be made and discussed.

This paper presents the results of an analytical study conducted to investigate airframe/engine interface dynamics, and the influence of rotor speed variations on the flight dynamics of the helicopter in hover, and to explore the potential benefits of using rotor states as additional feedback signals in the flight control system. The analytical investigation required the development of a parametric high-order helicopter hover model, which included heave/yaw body motion, the rotor speed degree of freedom, rotor blade motion in flapping and lead-lag, inflow dynamics, a drive train model with a flexible rotor shaft, and an engine/rpm governor. First, the model was used to gain insight into the engine/drive train/rotor system dynamics and to obtain an improved simple formula for easy estimation of the dominant first torsional mode, which is important in the dynamic integration of the engine and airframe system. Then, a linearized version of the model was used to investigate the effects of rotor speed variations and rotor state feedback on helicopter flight dynamics. Results show that, by including rotor speed variations, the effective vertical damping decreases significantly from that calculated with a constant speed assumption, thereby providing a better correlation with flight test data. Higher closed-loop bandwidths appear to be more readily achievable with rotor state feedback. The results also indicate that both aircraft and rotor flapping responses to gust disturbance are significantly attenuated when rotor state feedback is used.

Uh1

5% scale Blended-Wing-Body free-flight model in the Langley Full Scale Tunnel, building 640. People in the photographs include Sue Grafton (with model), Dan Vicroy (with model). Group photo: S. Grafton, D. Vicroy, D. Murri, C. Buttrill, J. Brandon, M. Croom, B. Jackson, C. Debro, G. Adams.

The Fluid Dynamics Branch's (ER42) at MSFC mission is to support NASA and other customers with discipline expertise to enable successful accomplishment of program/project goals. The branch is responsible for all aspects of the discipline of fluid dynamics, analysis and testing, applied to propulsion or propulsion-induced loads and environments, which includes the propellant delivery system, combustion devices, coupled systems, and launch and separation events. ER42 supports projects from design through development, and into anomaly and failure investigations. ER42 is committed to continually improving the state-of-its-practice to provide accurate, effective, and timely fluid dynamics assessments and in extending the state-of-the-art of the discipline.

No abstract available

The Supersonic Flight Dynamics Test is a full-scale flight test of aerodynamic decelerator technologies developed by the Low Density Supersonic Decelerator technology demonstration project. The purpose of the project is to develop and mature aerodynamic decelerator technologies for landing large-mass payloads on the surface of Mars. The technologies include a Supersonic Inflatable Aerodynamic Decelerator and supersonic parachutes. The first Supersonic Flight Dynamics Test occurred on June 28th, 2014 at the Pacific Missile Range Facility. The purpose of this test was to validate the test architecture for future tests. The flight was a success and, in addition, was able to acquire data on the aerodynamic performance of the supersonic inflatable decelerator. The Supersonic Disksail parachute developed a tear during deployment. The second flight test occurred on June 8th, 2015, and incorporated a Supersonic Ringsail parachute which was redesigned based on data from the first flight. Again, the inflatable decelerator functioned as predicted but the parachute was damaged during deployment. This paper describes the instrumentation, analysis techniques, and acquired flight test data utilized to reconstruct the vehicle trajectory, main motor thrust, atmosphere, and aerodynamics.

This paper is in the public domain in USA. Metadata comes from the CrossRef API, see full record in the source URL below.

This paper is in the public domain in USA. Metadata comes from the CrossRef API, see full record in the source URL below.

NAS 1.60:1700 Digitized from IA1177308-06-0000 , IA1177308-06-0001 .

NAS 1.60:1700 Digitized from IA1177308-06-0000 , IA1177308-06-0001 .

3 p. 23 cm

A modified F/A-18A undergoes wing torsion or twist testing in the Flight Dynamics Laboratory at NASA's Dryden Flight Research Center, Edwards, California. The tests simulate the amount of twisting that the wing might encounter during flight in the Active Aeroelastic Wing (AAW) flight research project at Dryden. During the tests, powerful hydraulic jacks attached to a fixture mounted on the tip of the F/A-18's left wing literally twist the wing to determine its flexibility, while numerous strain gauges and other devices record the stresses on the wing structure.

An overview of pre-flight aerodynamic models for the Low Density Supersonic Decelerator (LDSD) Supersonic Flight Dynamics Test (SFDT) campaign is presented, with comparisons to reconstructed flight data and discussion of model updates. The SFDT campaign objective is to test Supersonic Inflatable Aerodynamic Decelerator (SIAD) and large supersonic parachute technologies at high altitude Earth conditions relevant to entry, descent, and landing (EDL) at Mars. Nominal SIAD test conditions are attained by lifting a test vehicle (TV) to 36 km altitude with a large helium balloon, then accelerating the TV to Mach 4 and and 53 km altitude with a solid rocket motor. The first flight test (SFDT-1) delivered a 6 meter diameter robotic mission class decelerator (SIAD-R) to several seconds of flight on June 28, 2014, and was successful in demonstrating the SFDT flight system concept and SIAD-R. The trajectory was off-nominal, however, lofting to over 8 km higher than predicted in flight simulations. Comparisons between reconstructed flight data and aerodynamic models show that SIAD-R aerodynamic performance was in good agreement with pre-flight predictions. Similar comparisons of powered ascent phase aerodynamics show that the pre-flight model overpredicted TV pitch stability, leading to underprediction of trajectory peak altitude. Comparisons between pre-flight aerodynamic models and reconstructed flight data are shown, and changes to aerodynamic models using improved fidelity and knowledge gained from SFDT-1 are discussed.

A comprehensive set of motion equations for a multi-actuated flight vehicle is presented. The dynamics are derived from a vector approach that generalizes the classical linear perturbation equations for flexible launch vehicles into a coupled three-dimensional model. The effects of nozzle and aerosurface inertial coupling, sloshing propellant, and elasticity are incorporated without restrictions on the position, orientation, or number of model elements. The present formulation is well suited to matrix implementation for large-scale linear stability and sensitivity analysis and is also shown to be extensible to nonlinear time-domain simulation through the application of a special form of Lagrange s equations in quasi-coordinates. The model is validated through frequency-domain response comparison with a high-fidelity planar implementation.

Photographed on 05/17/2005. -- Group photographs of: 1) Flight Dyanmics Branch (FDB), 2) Dynamic Systems and Controls Branch (DSCB) and 3) Combined FDB and DSCB groups.