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At least 217 records · Page 12
All sourcesNASA NTRSDOE OSTI
NASA NTRS1987

Variable-Sweep Transition Flight Experiment (VSTFE): Stability code development and clean-up glove data analysis

The primary objective of the Variable Sweep Transition Flight Experiment (VSTFE) was to establish an improved swept wing transition criterion. The development of the Unified Stability System gave a way of quickly examining disturbance growth for a wide variety of laminar boundary layers. The disturbance growth traces shown are too scattered to define a transition criteria to replace the F-111 data band, which has been used successfully to design NLF gloves. Still, a careful review of the clean-up glove data may yield cases for which the transition location is known more accurately. Liquid crystal photographs of the clean-up glove show much spanwise variation in the transition front for some conditions, and this further complicates the analyses. Several high quality cases are needed in which the transition front is well defined and at a relatively constant chordwise station.

Rozendaal, Roger A.↗
NASA NTRS1990

Dielectric determination of the glass transition temperature (T sub g)

The objective is to determine the glass transition temperature of a polymer using a dielectric dissipation technique. A peak in the dissipation factor versus temperature curve is expected near the glass transition temperature T sub g. It should be noted that the glass transition is gradual rather than abrupt, so that the glass transition temperature T sub g is not clearly identifiable. In this case, the glass transition temperature is defined to be the temperature at the intersection point of the tangent lines to the dissipation factor versus temperature curve above and below the transition region, as illustrated.

Ries, Heidi R.↗
NASA NTRS1990

An investigation into the numerical prediction of boundary layer transition using the K.Y. Chien turbulence model

Assessments were made of the simulation capabilities of transition models developed at the University of Minnesota, as applied to the Launder-Sharma and Lam-Bremhorst two-equation turbulence models, and at The University of Texas at Austin, as applied to the K. Y. Chien two-equation turbulence model. A major shortcoming in the use of the basic K. Y. Chien turbulence model for low-Reynolds number flows was identified. The problem with the Chien model involved premature start of natural transition and a damped response as the simulation moved to fully turbulent flow at the end of transition. This is in contrast to the other two-equation turbulence models at comparable freestream turbulence conditions. The damping of the transition response of the Chien turbulence model leads to an inaccurate estimate of the start and end of transition for freestream turbulence levels greater than 1.0 percent and to difficulty in calculating proper model constants for the transition model.

Stephens, Craig A.↗
NASA NTRS1989

A transition detection study at Mach 1.5, 2.0, and 2.5 using a micro-thin hot-film system

A boundary-layer transition detection study was conducted in the NASA Langley unitary plan wind tunnel with an array of microthin hot films on a flat plate at Mach numbers 1.5, 2.0, and 2.5 and Reynolds numbers (1.0-4.5) x 10 to the 6th/ft. Transition locations were obtained online from the variation of normalized rms voltages from an array of hot-film sensors for both natural transition and a grit-induced wedge of turbulence. The effects of Mach number and Reynolds number on the location and length of the transition region for the two types of transition are presented from the online data. Also shown are the unit Reynolds number effects on transition Reynolds number, voltage-versus-time traces, spectra, and the data-acquisition system.

Johnson, Charles B.↗
NASA NTRS1990

Active-to-passive transition in the oxidation of silicon carbide and silicon nitride in air

The active-to-passive transition in the oxidation of SiC and Si3N4 was determined in a flowing air environment as a function of temperature and total pressure. The experimentally observed transition temperatures ranged from a low of 1347 C to a high of 1543 C for partial pressures of oxygen of 2.5 and 123.2 Pa, respectively. The SiC and Si3N4 samples had approximately the same transition point for a given pressure. In general, the higher the flow rate, the higher the transition temperature for a given pressure. The transitions for SiC measured in this study agree with previous data for the transition of SiC measured in pure oxygen at reduced pressures and in oxygen inert gas mixtures.

Vaughn, Wallace L.↗
NASA NTRS1990

Modelling the transitional boundary layer

Recent developments in the modelling of the transition zone in the boundary layer are reviewed (the zone being defined as extending from the station where intermittency begins to depart from zero to that where it is nearly unity). The value of using a new non-dimensional spot formation rate parameter, and the importance of allowing for so-called subtransitions within the transition zone, are both stressed. Models do reasonably well in constant pressure 2-dimensional flows, but in the presence of strong pressure gradients further improvements are needed. The linear combination approach works surprisingly well in most cases, but would not be so successful in situations where a purely laminar boundary layer would separate but a transitional one would not. Intermittency-weighted eddy viscosity methods do not predict peak surface parameters well without the introduction of an overshooting transition function whose connection with the spot theory of transition is obscure. Suggestions are made for further work that now appears necessary for developing improved models of the transition zone.

Narasimha, R.↗
NASA NTRS1991

An experimental comparison of nonswirling and swirling flow in a circular-to-rectangular transition duct

Circular-to-rectangular transition ducts are used as exhaust system components of aircraft with rectangular exhaust nozzles. Often, the incoming flow of these transition ducts includes a swirling velocity component remaining from the gas turbine engine. Previous transition duct studies have either not included inlet swirl or when inlet swirl was considered, only overall performance parameters were evaluated. Circular-to-rectangular transition duct flows with and without inlet swirl were explored in order to understand the effect of inlet swirl on the transition duct flow field and to provide detailed duct flow data for comparison with numerical code predictions. A method was devised to create a swirling, solid body rotational flow with minimal associated disturbances. Coefficients based on velocities and total and static pressures measured incross stream planes at four axial locations within the transition duct, along with surface static pressure measurements and surface oil film visualization, are presented for both nonswirling and swirling incoming flow. In both cases the inlet centerline Mach number was 0.35. The Reynolds number based on the inlet centerline velocity and duct inlet diameter was 1,547,000 for nonswirling and 1,366,000 for swirling flow. The maximum swirl angle was 15.6 deg. Two pair of counter-rotating side wall vortices appeared in the duct flow without inlet swirl. These vortices were absent in the swirling incoming flow cases.

Reichert, B. A.↗
NASA NTRS1991

Free-stream turbulence and concave curvature effects on heated, transitional boundary layers

An experimental investigation of the transition process on flat-plate and concave curved-wall boundary layers for various free-stream turbulence levels was performed. Results show that for transition of a flat-plate, the two forms of boundary layer behavior, identified as laminar-like and turbulent-like, cannot be thought of as separate Blasius and fully-turbulent profiles, respectively. Thus, simple transition models in which the desired quantity is assumed to be an average, weighted on intermittency, of the theoretical laminar and fully turbulent values is not expected to be successful. Deviation of the flow identified as laminar-like from theoretical laminar behavior is shown to be due to recovery after the passage of a turbulent spot, while deviation of the flow identified as turbulent-like from the full-turbulent values is thought to be due to incomplete establishment of the fully-turbulent power spectral distribution. Turbulent Prandtl numbers for the transitional flow, computed from measured shear stress, turbulent heat flux and mean velocity and temperature profiles, were less than unity. For the curved-wall case with low free-stream turbulence intensity, the existence of Gortler vortices on the concave wall within both laminar and turbulent flows was established using liquid crystal visualization and spanwise velocity and temperature traverses. Transition was found to occur via a vortex breakdown mode. The vortex wavelength was quite irregular in both the laminar and turbulent flows, but the vortices were stable in time and space. The upwash was found to be more unstable, with higher levels of u' and u'v', and lower skin friction coefficients and shape factors. Turbulent Prandtl numbers, measured using a triple-wire probe, were found to be near unity for all post-transitional profiles, indicating no gross violation of Reynolds analogy. No evidence of streamwise vortices was seen in the high turbulence intensity case.

Kim, J.↗
NASA NTRS1991

Review of boundary layer transition flight data on the Space Shuttle Orbiter

Both thermocouple data and flight-derived axial force coefficients have been used to determine the freestream flight conditions at which boundary-layer transition occurs on the windward surface of the Space Shuttle Orbiter during reentry. Boundary-layer-edge local flow conditions corresponding to transition at thermocouple locations on the Orbiter windward centerline have also been computed. Tables of the freestream conditions and the local flow conditions at transition are included. Transition occurred over a freestream Mach number range of 6.6 to 17.9 and Reynolds number range of (2.3-9.6) x 10 to the 6th. The roughness state of the Orbiter vehicles is also discussed and correlated to the occurrence of transition. Using the distributed roughness transition correlation of Bertin et al. (1982), a flight-derived roughness height of 0.061 inches was obtained. An average flight-derived roughness height of 0.103 inches for a discrete roughness element in the nose region was obtained using the correlation of Van Driest and Blumer (1968).

Bouslog, S. A.↗
NASA NTRS1992

Three-dimensional boundary-layer transition on a cone at Mach 3.5

A boundary-layer transition study on a sharp, 5 deg half-angle cone at various angles of attack was conducted at Mach 3.5. Transition data were obtained with and without significantly reduced freestream acoustic disturbance levels. A progressive downstream and upstream motion of the transition front on the windward and leeward rays, respectively, of the cone with angle of attack was observed for the high noise level data in agreement with data trends obtained in conventional ('noisy') wind tunnels. However, the downstream movement was not observed to the same degree for the low noise level data in the present study. Transition believed to be crossflow dominated was found to be less receptive to freestream acoustic disturbances than first-mode (Tollmien-Schlichting) dominated transition. The previously-developed crossflow transition Reynolds number criterion was found to be inadequate for the current case. An improved criterion is offered, which includes compressibility and flow-geometry effects.

King, R. A.↗
NASA NTRS1991

Late-time cosmological phase transitions

It is shown that the potential galaxy formation and large scale structure problems of objects existing at high redshifts (Z approx. greater than 5), structures existing on scales of 100 M pc as well as velocity flows on such scales, and minimal microwave anisotropies ((Delta)T/T) (approx. less than 10(exp -5)) can be solved if the seeds needed to generate structure form in a vacuum phase transition after decoupling. It is argued that the basic physics of such a phase transition is no more exotic than that utilized in the more traditional GUT scale phase transitions, and that, just as in the GUT case, significant random Gaussian fluctuations and/or topological defects can form. Scale lengths of approx. 100 M pc for large scale structure as well as approx. 1 M pc for galaxy formation occur naturally. Possible support for new physics that might be associated with such a late-time transition comes from the preliminary results of the SAGE solar neutrino experiment, implying neutrino flavor mixing with values similar to those required for a late-time transition. It is also noted that a see-saw model for the neutrino masses might also imply a tau neutrino mass that is an ideal hot dark matter candidate. However, in general either hot or cold dark matter can be consistent with a late-time transition.

Schramm, David N.↗
NASA NTRS1938

Determination of Boundary-Layer Transition on Three Symmetrical Airfoils in the NACA Full-Scale Wind Tunnel

For the purpose of studying the transition from laminar to turbulent flow, boundary-layer measurements were made in the NACA full-scale wind tunnel on three symmetrical airfoils of NACA 0009, 0012, and 0018 sections. The effects of variations in lift coefficient, Reynolds number, and airfoil thickness on transition were investigated. Air speed in the boundary layer was measured by total-head tubes and by hot wires; a comparison of transition as indicated by the two techniques was obtained. The results indicate no unique value of Reynolds number for the transition, whether the Reynolds number is based upon the distance along the chord or upon the thickness of the boundary layer at the transition point. In general, the transition is not abrupt and occurs in a region that varies in length as a function of the test conditions.

Silverstein, Abe↗
NASA NTRS1992

Linear stability theory and three-dimensional boundary layer transition

The viewgraphs and discussion of linear stability theory and three dimensional boundary layer transition are provided. The ability to predict, using analytical tools, the location of boundary layer transition over aircraft-type configurations is of great importance to designers interested in laminar flow control (LFC). The e(sup N) method has proven to be fairly effective in predicting, in a consistent manner, the location of the onset of transition for simple geometries in low disturbance environments. This method provides a correlation between the most amplified single normal mode and the experimental location of the onset of transition. Studies indicate that values of N between 8 and 10 correlate well with the onset of transition. For most previous calculations, the mean flows were restricted to two-dimensional or axisymmetric cases, or have employed simple three-dimensional mean flows (e.g., rotating disk, infinite swept wing, or tapered swept wing with straight isobars). Unfortunately, for flows over general wing configurations, and for nearly all flows over fuselage-type bodies at incidence, the analysis of fully three-dimensional flow fields is required. Results obtained for the linear stability of fully three-dimensional boundary layers formed over both wing and fuselage-type geometries, and for both high and low speed flows are discussed. When possible, transition estimates form the e(sup N) method are compared to experimentally determined locations. The stability calculations are made using a modified version of the linear stability code COSAL. Mean flows were computed using both Navier Stokes and boundary-layer codes.

Spall, Robert E.↗
NASA NTRS1996

Structure Identification Within a Transitioning Swept-Wing Boundary Layer

Extensive measurements are made in a transitioning swept-wing boundary layer using hot-film, hot-wire and cross-wire anemometry. The crossflow-dominated flow contains stationary vortices that breakdown near mid-chord. The most amplified vortex wavelength is forced by the use of artificial roughness elements near the leading edge. Two-component velocity and spanwise surface shear-stress correlation measurements are made at two constant chord locations, before and after transition. Streamwise surface shear stresses are also measured through the entire transition region. Correlation techniques are used to identify stationary structures in the laminar regime and coherent structures in the turbulent regime. Basic techniques include observation of the spatial correlations and the spatially distributed auto-spectra. The primary and secondary instability mechanisms are identified in the spectra in all measured fields. The primary mechanism is seen to grow, cause transition and produce large-scale turbulence. The secondary mechanism grows through the entire transition region and produces the small-scale turbulence. Advanced techniques use Linear Stochastic Estimation (LSE) and Proper Orthogonal Decomposition (POD) to identify the spatio-temporal evolutions of structures in the boundary layer. LSE is used to estimate the instantaneous velocity fields using temporal data from just two spatial locations and the spatial correlations. Reference locations are selected using maximum RMS values to provide the best available estimates. POD is used to objectively determine modes characteristic of the measured flow based on energy. The stationary vortices are identified in the first laminar modes of each velocity component and shear component. Experimental evidence suggests that neighboring vortices interact and produce large coherent structures with spanwise periodicity at double the stationary vortex wavelength. An objective transition region detection method is developed using streamwise spatial POD solutions which isolate the growth of the primary and secondary instability mechanisms in the first and second modes, respectively. Temporal evolutions of dominant POD modes in all measured fields are calculated. These scalar POD coefficients contain the integrated characteristics of the entire field, greatly reducing the amount of data to characterize the instantaneous field. These modes may then be used to train future flow control algorithms based on neural networks.

Chapman, Keith↗
NASA NTRS1997

A New Approach for Determining Onset of Transition

The final report consists of three papers which outline and demonstrate the new method for determining transition onset. The procedure developed under this grant requires specification of the instability mechanism, i.e., Tollmien-Schlichting or crossflow, that leads to transition. The attached papers are entitled: 'An Alternative to the e(sup n) Method for Determining Onset of Transition', 'Transition Model for Swept Wing Flows', and 'A Transition Closure Model for Predicting Transition Onset'.

Hassan, H. A.↗
NASA NTRS1998

Direct Numerical Simulations of Boundary Layer Transition on a Flat Plate

In recent years the techniques of computational fluid dynamics (CFD) have been used to compute flows associated with geometrically complex configurations. However, success in terms of accuracy and reliability has been limited to cases where the effects of turbulence and transition could be modeled in a straightforward manner. Even in simple flows, the accurate computation of skin friction and heat transfer using existing turbulence models has proved to be a difficult task, one that has required extensive fine-tuning of the turbulence models used. In more complex flows (for example, in turbomachinery flows in which vortices and wakes impinge on airfoil surfaces causing periodic transitions from laminar to turbulent flow) the development of a model that accounts for all scales of turbulence and predicts the onset of transition is an extremely difficult task. Fortunately, current trends in computing suggest that it may be possible to perform direct simulations of turbulence and transition at moderate Reynolds numbers in some complex cases in the near future. This presentation will focus on direct simulations of transition and turbulence using high-order accurate finite-difference methods. The advantage of the finite-difference approach over spectral methods is that complex geometries can be treated in a straightforward manner. Additionally, finite-difference techniques are the prevailing methods in existing application codes. An application of high-order-accurate finite-difference methods to direct simulations of transition and turbulence in a spatially evolving boundary layer subjected to high levels of freestream turbulence will be presented.

Rai, Man Mohan↗
NASA NTRS1998

Transition in Turbine Flows

We have further developed our capabilities to analyze transition in turbine boundary layers from first principles by integrating the nonlinear parabolized stability equations (PSE) with improved initial and boundary conditions. With modified iteration schemes, we are able to proceed deeper into the transition region where skin friction coefficient and heat transfer coefficient significantly increase. Initial and boundary conditions at elevated turbulence levels can be derived by receptivity analysis. Test runs for ERCOFTAC test case T3A at 2.4\% turbulence level provide results in good agreement with the experimental data. The sharper minimum of the skin coefficient also shown by DNS results is likely due to the missing intermittency. The method has been applied to various experimentally studied turbine blades (UTRC, VKI, Zierke, Langston, Hippensteele, and others). The PSE results, though physically reasonable, do not agree as well as expected with the experimental findings. We have, therefore, performed an extensive search for the reasons of the seemingly systematic deviations. A first source of uncertainty has been found in the often insufficient documentation of the experiments (e.g. on blockage by end-wall boundary layers). However, variation of the relevant parameters does not lead to more satisfactory agreement. A second reason has been found in the "standard procedure" which considers a 2D flow at midspan and uses a panel code and subsequent boundary-layer code to obtain the laminar basic flow for the transition analysis. Comparison with the pressure distribution obtained with a 3D design code (RVC3D) shows significant three-dimensionality of the flow (e.g. in the UTRC experiments). The spanwise variation has been neglected in our original PSE code. To overcome this problem, we have developed the PSE/3D for fully 3D boundary layers to account for streamwise and spanwise variations. Since the design code does not provide the boundary-layer flow with sufficient resolution, we have generated the Euler solution and employed a 3D boundary-layer code to obtain the viscous basic flow. Although only the linear stability level of PSE/3D has been implemented so far, the discrepancies with the experiments change but do not disappear. We still find deviations between the computed and experimental variations of C(sub f), and St along the blade for laminar flow. The main reason can be seen by comparing the solution of the boundary-layer code with the viscous results of the design code. The conventional boundary-layer solution exhibits an asymptotic behavior appropriate in external aerodynamics but does not match the steep gradients of the inviscid flow through the passage and consequently provides biased results for C(sub f), and St. An attempt is currently being made to correct this deficiency. Before attempting to perform the transition analysis for the viscous flow provided by the design code, we have analyzed the implementation and "best possible" results. Code and results exhibit flaws that may negatively affect the design and are intolerable for transition analysis. Therefore, we have decided to develop a new code to obtain a reliable basis for stability and transition studies. We expect to report improved results by the time of the meeting.

Herbert, Thorwald↗
NASA NTRS1998

Mach Number Effects on Turbine Blade Transition Length Prediction

The effect of a Mach number correction on a model for predicting the length of transition was investigated. The transition length decreases as the turbulent spot production rate increases. Much of the data for predicting the spot production rate comes from low speed flow experiments. Recent data and analysis showed that the spot production rate is affected by Mach number. The degree of agreement between analysis and data for turbine blade heat transfer without film cooling is strongly dependent of accurately predicting the length of transition. Consequently, turbine blade heat transfer data sets were used to validate a transition length turbulence model. A method for modifying models for the length of transition to account for Mach number effects is presented. The modification was made to two transition length models. The modified models were incorporated into the two-dimensional Navier-Stokes code, RVCQ3D. Comparisons were made between predicted and measured midspan surface heat transfer for stator and rotor turbine blades. The results showed that accounting for Mach number effects significantly improved the agreement with the experimental data.

Boyle, R. J.↗