Engineering topics
Publications and source records attributed to Kerrebrock, J. L..
Appendices A to F present the theory behind the TASOPT methodology and code. Appendix A describes the bulk of the formulation, while Appendices B to F develop the major sub-models for the engine, fuselage drag, BLI accounting, etc.
MIT, Aerodyne Research, Aurora Flight Sciences, and Pratt & Whitney have collaborated to address NASA s desire to pursue revolutionary conceptual designs for a subsonic commercial transport that could enter service in the 2035 timeframe. The MIT team brings together multidisciplinary expertise and cutting-edge technologies to determine, in a rigorous and objective manner, the potential for improvements in noise, emissions, and performance for subsonic fixed wing transport aircraft. The collaboration incorporates assessment of the trade space in aerodynamics, propulsion, operations, and structures to ensure that the full spectrum of improvements is identified. Although the analysis focuses on these key areas, the team has taken a system-level approach to find the integrated solutions that offer the best balance in performance enhancements. Based on the trade space analyses and system-level assessment, two aircraft have been identified and carried through conceptual design to show both the in-depth engineering that underpins the benefits envisioned and also the technology paths that need to be followed to enable, within the next 25 years, the development of aircraft three generations ahead in capabilities from those flying today.
Under NASA Glenn Research Center sponsorship, MIT has developed the concept of micromachined, bipropellant, liquid rocket engines. This is potentially a breakthrough technology changing the cost-performance tradeoffs for small propulsion systems, enabling new applications, and redefining the meaning of the term low-cost-access-to-space. With this NASA support, a liquid-cooled, gaseous propellant version of the thrust chamber and nozzle was designed, built, and tested as a first step. DARPA is currently funding MIT to demonstrate turbopumps and controls. The work performed herein was the second year of a proposed three-year effort to develop the technology and demonstrate very high power density, regeneratively cooled, liquid bipropellant rocket engine thrust chamber and nozzles. When combined with the DARPA turbopumps and controls, this work would enable the design and demonstration of a complete rocket propulsion system. The original MIT-NASA concept used liquid oxygen-ethanol propellants. The military applications important to DARPA imply that storable liquid propellants are needed. Thus, MIT examined various storable propellant combinations including N2O4 and hydrazine, and H2O2 and various hydrocarbons. The latter are preferred since they do not have the toxicity of N2O4 and hydrazine. In reflection of the newfound interest in H2O2, it is once again in production and available commercially. A critical issue for the microrocket engine concept is cooling of the walls in a regenerative design. This is even more important at microscale than for large engines due to cube-square scaling considerations. Furthermore, the coolant behavior of rocket propellants has not been characterized at microscale. Therefore, MIT designed and constructed an apparatus expressly for this purpose. The report details measurements of two candidate microrocket fuels, JP-7 and JP-10.
A three-dimensional model of the stability of the particle-bed reactor is presented, in which the fluid has mobility in three dimensions. The model accurately represents the stability at low Re numbers as well as the effects of the cold and hot frits and of the heat conduction and radiation in the particle bed. The model can be easily extended to apply to the cylindrical geometry of particle-bed reactors. Exemplary calculations are carried out, showing that a particle bed without a cold frit would be subject to instability if operated at the high-temperature ratios used for nuclear rockets and at power densities below about 4 MW/l; since the desired power density for such a reactor is about 40 MW/l, the operation at design exit temperature but at reduced power could be hazardous. Calculations show however that it might be possible to remove the instability problem by appropriate combinations of cold and hot frits.
The present consideration of rotation effects on heat transfer in a radially exhausted, impingement-cooled turbine blade model gives attention to experimental results for Reynolds and Rossby numbers and blade/coolant temperature ratio values that are representative of small gas turbine engines. On the basis of a model that encompasses the effects of Coriolis force and buoyancy on heat transfer, bouyancy is identified as the cause of an average Nusselt number that is 20-30 percent lower than expected from previous nonrotating data. A heuristic model is proposed which predicts that the impingement jets nearest the blade roots should deflect inward, due to a centripetal force generated by their tangential velocity counter to the blade motion. Potentially serious thermal stresses must be anticipated from rotation effects in the course of blade design.
The effects of rotation on impingement cooling of turbine blades were studied experimentally as a specialized facility at M.I.T. A foil heated resistively was cooled by a jet flow on one side and temperature monitored on the other. Rotating the blade limits the heat transfer path to conduction through the support structure and radiation. IR radiometry furnishes the temperature distributions on the chamber wall, permitting the internal heat transfer coefficient to be measured. The heat transfer efficiency has been found to fall as much as 30 percent as rotational speed increases. The conditions observed confirm the significance of rotational effects, particularly with regard to potential early blade failure.
The time resolved flowfield in a low aspect ratio transonic compressor stage has been studied using a high frequency response sphere probe with a bandpass of D.C. to 20 kHz in a blowdown compressor facility. Averaged over the compressor annulus, the data agree well with those measured with standard pitot type probes on the same stage in a conventional test rig. Not all the spanwise distributions agree, however. These differences are explained as errors in the pitot probe readings due to fluctuations in the flow. The experimental data are compared to the results of a three-dimensional inviscid Euler calculation.
A spherical probe having five surface mounted silicon pressure sensors is described. The probe is 0.2 inches in diameter yielding a frequency response above 30 kHz. A scheme is described for retrieving from the five pressure signals the stagnation and static pressures, Mach number and flow angles in two planes. The calibrations forming the base for this procedure, obtained from steady state tests, are given in the form of polynomials in the two flow angles. A method for compensating the silicon pressure sensors for thermal effects is given. Typical data obtained in the MIT Blowdown Compressor are presented in raw form and after reduction to flow angles, Mach number and stagnation pressure.
The aeroelastic behavior of a transonic compressor rotor operated in the MIT Blowdown Compressor Facility has been examined by means of piezoelectric motion sensors at the base of each of the 23 blades. Excitation has been observed due to rotating stall, due to an incipient flutter, and due to the facility startup transient. A method has been found for determining the aerodynamic damping force by modal analysis of the blade motion. Application of this technique to the example of excitation by rotating stall has led to the conclusions that the blade loading decreases in the stall cell, and that the damping force on the blades in the clean flow is in phase with blade velocity but opposite it in sign, leading to a logarithmic decrement of 0.2. This method of force derivation has quite general applicability as it requires only blade motion data such as are routinely acquired with strain gages. It is argued that models are needed for aerodynamic damping which focus on the effects of near neighbors of a given blade, since flutter often results in large response of isolated blades or small groups of blades.
The inadequacy of axisymmetric theory in modeling the structure of the shock system in a transonic compressor rotor is shown with a simple model. A more accurate model for the flow near the sonic radius is proposed, which is supported by some flow visualization data and three-dimensional computations.
Some results of a comprehensive study of the flow through a transonic compressor rotor are presented. The rotor produces a pressure ratio of 1.6 at a tip Mach number of 1.2, has a cylindrical casing, sloped hub, and MCA blade sections. The gas density has been measured in the rotor by gas fluorescence, which reveals details of the shock and boundary layer structure as well as density in the passages. Time resolved measurements of the flow just downstream of the rotor give the blade wake structure, which is related to the features shown by the fluorescence. Of particular interest is a complex three-dimensional separation near the sonic radius. The flow through the rotor has also been computed in three dimensions by a time-marching technique, and the results of this computation are compared to those found by flow visualization.
A detailed time resolved study of the flow field upstream and downstream of a high work transonic compressor rotor shows that the flow field is dominated by the downstream evolution of the viscous flow shed from the rotor blades under the influence of the strong mean swirl. The dominant periodicity in the flow changes from blade passing to 1.4 times blade passing within one chord from the blade row. A possible explanation is that the wakes evolve to a shear eigenmode of the swirling flow, as suggested by perturbation theory. Another possibility is a 'propagating stall' of 16 cells, but the rotor operated near its design point. Treatments of 'turbulence' in turbomachines should account for such phenomena, which originate in the strong mean swirl.
The three dimensional unsteady flow field behind a transonic compressor rotor with a design pressure ratio of 1.6 at a tip Mach number of 1.2 has been resolved on the blade passing time scale. Quantities determined were total and static pressures, tangential flow angle, and radial flow angle. The spatial and temporal resolution achieved was sufficient to determine velocity components inside individual blade wakes and in the surrounding flow. From these measurements the flow structure is described at stations immediately behind the rotor and one chord downstream. Some dominant features of the flow just behind the rotor are large radial velocity components, large static pressure fluctuations near the blade wakes, and definite unsteadiness (in rotor coordinates) of the wakes. The wake behavior one chord downstream is described in terms of the effect of the strong mean swirl on the behavior of shear disturbances. In the outer portion of the annulus, where the mean flow approximates a solid body rotation, a strong, persistent oscillatory flow is found with 16 periods in the circumference as roughly predicted by theory. In the inner portion of the annulus the disturbances attenuate axially.
A blowdown compressor test facility has been developed which allows time-resolved aerodynamic testing of full-scale transonic compressor rotors at low cost. The rotor is brought to speed in vacuum, a diaphragm is opened, and the test gas is allowed to flow for a time of the order of 0.1 sec, during which the rotor is driven by its own inertia. Both 'steady-state' performance evaluation and detailed time resolution of the flow on the blade-passing time scale have been demonstrated for a two-ft-diam transonic rotor with a tangential Mach number of 1.2 and a nominal pressure ratio of 1.6. The steady-state performance as determined in the experiments includes an efficiency of 0.92 and a pressure ratio of 1.55 at design speed. The time-resolved measurements include the combination tone structure in the upstream flow field, resolved both axially and radially, and the wake structure downstream of the rotor, also resolved both radially and axially.
Acoustic instability driven by thermal radiation absorption in extremely high temperature gases
Rotor wakes intrastator transport effects on high Mach number axial flow compressors performance, considering stagnation temperature profile and rotor blade loss factor
Nonequlibrium plasma stability, conductivity and Hall parameter influenced by current flow parallel to magnetic field, discussing Ar-Cs and He-Cs plasma data
Design, operation, and evaluation of large, segmented electrode nonequilibrium magnetohydrodynamic generator