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Predicted and flight-test drag on the C-5A and the C-141 are correlated. Equivalent rigid flight-test profile drag and a rigid estimate based on wind tunnel data are also correlated. Correlations for the National Transonic Facility are included.
Data obtained through very high angles of attack from a large-scale, subsonic wind-tunnel test of a close-coupled canard-delta-wing fighter model are analyzed. The canard delays wing leading-edge vortex breakdown, even for angles of attack at which the canard is completely stalled. A vortex-lattice method was applied which gave good predictions of lift and pitching moment up to an angle of attack of about 20 deg, where vortex-breakdown effects on performance become significant. Pitch-control inputs generally retain full effectiveness up to the angle of attack of maximum lift, beyond which, effectiveness drops off rapidly. A high-angle-of-attack prediction method gives good estimates of lift and drag for the completely stalled aircraft. Roll asymmetry observed at zero sideslip is apparently caused by an asymmetry in the model support structure.
An evaluation is made of six diverse examples of nonlinear treatments of slender body theory for the prediction of missile aerodynamic behavior. The cases in question are the application of area rule to store carriage design in the drag rise region, the estimation of destabilizing pitching moments associated with transonic projectiles, the pressure loadings on elliptical missile airframes, nonlinear control characteristics, roll control effectiveness in canard missile configurations, and novel approaches for vortex flow modeling.
The outboard planforms of wings have been found to be of prime importance in studies of induced drag reduction. This conclusion is based on an experimental and theoretical study of the aerodynamic characteristics of planar and nonplanar outboard wing forms. Six different configurations; baseline rectangular, planar sheared, sheared with dihedral, sheared with anhedral, rising arc, and drooping arc were investigated for two different spans. Span efficiencies as much as 20 percent greater than baseline can be realized with nonplanar wing forms. Optimization studies show that this advantage can be achieved along with a bending moment benefit. Parasite drag and lateral stability estimations were not included in the analysis.
Future NASA Earth Observing System (EOS) Spacecraft will make measurements of the earth's clouds, oceans, atmosphere, land and radiation balance. These EOS Spacecraft will be part of the NASA Mission to Planet Earth. This paper specifically addresses the EOS AM Spacecraft, referred to as 'AM' because it has a sun-synchronous orbit with a 10:30 AM descending node. This paper describes the EOS AM Spacecraft mission orbit requirements, orbit determination, orbit control, and navigation system impact on earth based pointing. The EOS AM Spacecraft will be the first spacecraft to use the TDRSS Onboard Navigation System (TONS) as the primary means of navigation. TONS flight software will process one-way forward Doppler measurements taken during scheduled TDRSS contacts. An extended Kalman filter will estimate spacecraft position, velocity, drag coefficient correction, and ultrastable master oscillator frequency bias and drift. The TONS baseline algorithms, software, and hardware implementation are described in this paper. TONS integration into the EOS AM Spacecraft Guidance, Navigation, and Control (GN&C) System; TONS assisted onboard time maintenance; and the TONS Ground Support System (TGSS) are also addressed.
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The Marshall engineering thermosphere model was specified by NASA to be used in the design, development and testing phases of the International Space Station (ISS). The mass density is the atmospheric parameter which most affects the ISS. Under simplifying assumptions, the critical ambient neutral density required to produce one micro-g on the ISS is estimated using an atmospheric drag acceleration equation. Examples are presented for the critical density versus altitude, and for the critical density that is exceeded at least once a month and once per orbit during periods of low and high solar activity. An analysis of the ISS orbital decay is presented.
Recently acquired experimental data for the DLR-F6 wing-body transonic transport con figuration from the National Transonic Facility (NTF) are compared with the database of computational fluid dynamics (CFD) predictions generated for the Third AIAA CFD Drag Prediction Workshop (DPW-III). The NTF data were collected after the DPW-III, which was conducted with blind test cases. These data include both absolute drag levels and increments associated with this wing-body geometry. The baseline DLR-F6 wing-body geometry is also augmented with a side-of-body fairing which eliminates the flow separation in this juncture region. A comparison between computed and experimentally observed sizes of the side-of-body flow-separation bubble is included. The CFD results for the drag polars and separation bubble sizes are computed on grids which represent current engineering best practices for drag predictions. In addition to these data, a more rigorous attempt to predict absolute drag at the design point is provided. Here, a series of three grid densities are utilized to establish an asymptotic trend of computed drag with respect to grid convergence. This trend is then extrapolated to estimate a grid-converged absolute drag level.
The aerodynamic performance of rotorcraft designed for heavy-lift and high-speed cruise is examined. Configurations considered include the tiltrotor, the compound helicopter, and the lift-offset rotor. Design conditions are hover and 250-350 knot cruise, at 5k/ISA+20oC (civil) or 4k/95oF (military); with cruise conditions at 4000 or 30,000 ft. The performance was calculated using the comprehensive analysis CAMRAD II, emphasizing rotor optimization and performance, including wing-rotor interference. Aircraft performance was calculated using estimates of the aircraft drag and auxiliary propulsion efficiency. The performance metric is total power, in terms of equivalent aircraft lift-to-drag ratio L/D = WV/P for cruise, and figure of merit for hover.
Controlling discretization error is a remaining challenge for computational fluid dynamics simulation. Grid adaptation is applied to reduce estimated discretization error in drag or pressure integral output functions. To enable application to high O(10(exp 7)) Reynolds number turbulent flows, a hybrid approach is utilized that freezes the near-wall boundary layer grids and adapts the grid away from the no slip boundaries. The hybrid approach is not applicable to problems with under resolved initial boundary layer grids, but is a powerful technique for problems with important off-body anisotropic features. Supersonic nozzle plume, turbulent flat plate, and shock-boundary layer interaction examples are presented with comparisons to experimental measurements of pressure and velocity. Adapted grids are produced that resolve off-body features in locations that are not known a priori.
JPL is NASA's prime center for deep space missions. In response to the need to reduce the cost and time to complete early concept studies and proposals JPL created the first concurrent engineering team in the aerospace industry: Team X. Started in 1995, Team X has carried out over 800 studies, dramatically reducing the time and cost involved, and has been the model for other concurrent engineering teams both within NASA and throughout the larger aerospace community. Since its inception, the software backbone of this highly successful design team - engaged in examining some of NASA's cutting edge concepts - has been the unassuming spreadsheet. Over the years the Team X spreadsheet-based tools have evolved from simple standalone engineering models into a networked spreadsheet intensive system with real time parameter updating. Recent new capabilities include stochastic cost estimation and a graphical drag and drop block diagram that automatically populates the related spreadsheet parameters of cost, mass and power. This paper describes how the spreadsheet functions within Team X: its history, architecture, current capabilities, enabling strengths and persistent weaknesses. In addition, the verification methods and institutional oversight that have evolved as the Team X products became increasingly critical to Laboratory success are also discussed.
Empirical orthogonal function modeling is explained and applied to identify compact discrete-time nonlinear unsteady aerodynamic models from data generated by an unsteady three-dimensional compressible Navier-Stokes flow solver for an airfoil undergoing various pitching motions. Model structures, model parameter estimates, and model parameter uncertainty estimates for nondimensional lift, drag, and pitching moment coefficient models were determined autonomously and directly from the data. Prediction tests using data that were not used in the modeling process showed that the identified models exhibited excellent prediction capability, which is a strong indicator of an accurate model.
Four 0.01643-scale equivalent-body-of-revolution models, designed to aid in the evaluation of the relative merits of various degrees of redesign of the existing (1955) Convair F-102 airplane, were launched from the helium gun at Wallops Island, Va., to determine their zero-lift drag at Mach numbers from 0.8 to 1.3. The data are presented with only sufficient analysis to validate their general subsonic level. Estimated values of the friction drag are presented at all Mach numbers to allow a comparison of the pressure drag values alone.
Passive turbulent drag reduction techniques are of interest as a cost effective means to improve air vehicle fuel consumption. In the past, rigid surface waves slanted at an angle from the streamwise direction were deemed ineffective to reduce skin friction drag due to the pressure drag that they generate. A recent analysis seeking similarities to the spanwise shear stress generated by spatial Stokes layers suggested that there may be a range of wavelength, amplitude, and orientation in which the wavy surface would reduce turbulent drag. The present work explores, by experiments and Direct Numerical Simulations (DNS), the effect of swept wavy surfaces on skin friction and pressure drag. Plates with shallow and deep wave patterns were rapid-prototyped and tested using a drag balance in the 7x11 inch Low-Speed Wind Tunnel at the NASA LaRC Research Center. The measured drag o set between the wavy plates and the reference at plate is found to be within the experimental repeatability limit. Oil vapor flow measurements indicate a mean spanwise flow over the deep waves. The turbulent flow in channels with at walls, swept wavy walls and spatial Stokes spanwise velocity forcing was simulated at a friction Reynolds number of two hundred. The time-averaged and dynamic turbulent flow characteristics of the three channel types are compared. The drag obtained for the channel with shallow waves is slightly larger than for the at channel, within the range of the experiments. In the case of the large waves, the simulation over predicts the drag. The shortcomings of the Stokes layer analogy model for the estimation of the spanwise shear stress and drag are discussed.
A procedure for estimating fuel burned based on actual flight track data, and drag and fuel-flow models is described. The procedure consists of estimating aircraft and wind states, lift, drag and thrust. Fuel-flow for jet aircraft is determined in terms of thrust, true airspeed and altitude as prescribed by the Base of Aircraft Data fuel-flow model. This paper provides a theoretical foundation for computing fuel-flow with most of the information derived from actual flight data. The procedure does not require an explicit model of thrust and calibrated airspeed/Mach profile which are typically needed for trajectory synthesis. To validate the fuel computation method, flight test data provided by the Federal Aviation Administration were processed. Results from this method show that fuel consumed can be estimated within 1% of the actual fuel consumed in the flight test. Next, fuel consumption was estimated with simplified lift and thrust models. Results show negligible difference with respect to the full model without simplifications. An iterative takeoff weight estimation procedure is described for estimating fuel consumption, when takeoff weight is unavailable, and for establishing fuel consumption uncertainty bounds. Finally, the suitability of using radar-based position information for fuel estimation is examined. It is shown that fuel usage could be estimated within 5.4% of the actual value using positions reported in the Airline Situation Display to Industry data with simplified models and iterative takeoff weight computation.
The worldwide effort to design and commission floating offshore wind turbines (FOWT) is motivating the need for reliable numerical models that adequately represent their physical behavior under realistic sea states. However, properly representing the hydrodynamic quadratic damping for FOWT remains uncertain, because of its dependency on the choice of drag coefficients (dimensionless or not). It is hypothesized that the limited degree of freedom (DoF) drag coefficient formulation that uses only translational drag coefficients causes mischaracterization of the rotational DoF drag, leading to underestimation of FOWT global loads, such as tower base fore-aft shear. To address these hydrodynamic modeling uncertainties, different quadratic drag models implemented in the open-source mid-fidelity simulation tool, OpenFAST, were investigated and compared with the experimental data from the Offshore Code Comparison Collaboration, Continued, with Correlation (OC5) project. The tower base fore-aft shear and up-wave mooring line tension were compared under an irregular wave loading condition to demonstrate the effects of the different damping models. Two types of hydrodynamic quadratic drag formulations were considered: (1) member-based dimensionless drag coefficients applied only at the translational DoF (namely limited-DoF drag model) and (2) quadratic drag matrix model (in dimensional form). Based on the results, the former consistently underestimated the 95th percentile peak loads and spectral responses when compared to the OC5 experimental data. In contrast, the drag matrix models reduced errors in estimates of the tower base shear peak load by 7–10% compared to the limited-DoF drag model. The underestimation in the tower base fore-aft shear was thus inferred be related to mischaracterization of the rotational pitch drag and the heave motion/drag by the limited-DoF model.
This paper documents a series of estimates of the orbital lifetime of the SEDS-2 flight configuration made prior to the mission. These estimates were made with program LTIME, which has been in use at MSFC for a number of years. Because of the unusual configuration of upper-stage/tether/endmass flown on this mission, and the type of assumptions and inputs used in LTIME, the effective area used in the drag calculation had to be estimated in an unusual way. The final pre-flight predicted lifetime was 28.35 days. In the actual flight, the tether was cut approximately 5 days into the mission. The instrumented endmass plus about 12 km of tether rapidly reentered the atmosphere, and the Delta II Second Stage plus the remaining 8 km of tether reentered on mission day 60. Tracking data was used to reconstruct reentry sequences for the two parts of the configuration after the cut. The predicted lifetimes for the endmass plus tether-fragment were in the range of 0.2 to 2.8 days, depending on the perigee altitudes assumed. The predicted lifetime of the upper-stage plus tether was 56.4 days, which corresponds to reentry on mission day 61, in good agreement with the actual reentry on day 60.
An aircraft configuration, previously conceived as a means to achieve favorable aerodynamic stability characteristics., high lift-drag ratio, and low heating rates at high supersonic speeds., was modified in an attempt to increase further the lift-drag ratio without adversely affecting the other desirable characteristics. The original configuration consisted of three identical triangular wing panels symmetrically disposed about an ogive-cylinder body equal in length to the root chord of the panels. This configuration was modified by altering the angular disposition of the wing panels, by reducing the area of the panel forming the vertical fin, and by reshaping the body to produce interference lift. Six-component force and moment tests of the modified configuration at combined angles of attack and sideslip were made at a Mach number of 3.3 and a Reynolds number of 5.46 million. A maximum lift-drag ratio of 6.65 (excluding base drag) was measured at a lift coefficient of 0.100 and an angle of attack of 3.60. The lift-drag ratio remained greater than 3 up to lift coefficient of 0.35. Performance estimates, which predicted a maximum lift-drag ratio for the modified configuration 27 percent greater than that of the original configuration, agreed well with experiment. The modified configuration exhibited favorable static stability characteristics within the test range. Longitudinal and directional centers of pressure were slightly aft of the respective centroids of projected plan-form and side area.