Static aeroelastic analysis and tailoring of missile control fins

A concept for enhancing the design of control fins for supersonic tactical missiles is described. The concept makes use of aeroelastic tailoring to create fin designs (for given planforms) that limit the variations in hinge moments that can occur during maneuvers involving high load factors and high angles of attack. It combines supersonic nonlinear aerodynamic load calculations with finite-element structural modeling, static and dynamic structural analysis, and optimization. The problem definition is illustrated. The fin is at least partly made up of a composite material. The layup is fixed, and the orientations of the material principal axes are allowed to vary; these are the design variables. The objective is the magnitude of the difference between the chordwise location of the center of pressure and its desired location, calculated for a given flight condition. Three types of constraints can be imposed: upper bounds on static displacements for a given set of load conditions, lower bounds on specified natural frequencies, and upper bounds on the critical flutter damping parameter at a given set of flight speeds and altitudes. The idea is to seek designs that reduce variations in hinge moments that would otherwise occur. The block diagram describes the operation of the computer program that accomplishes these tasks. There is an option for a single analysis in addition to the optimization

Calculation of aerodynamic characteristics of STOL aircraft

Method predicts lift and pitching moment characteristics of STOL aircraft with externally-blown, jet-augmented wing-flap combinations using potential-flow approach which involves combination of two flow models. Method can accommodate multiple engines per wing panel and part-span flaps

Computation of aerodynamic interference between lifting surfaces and lift- and cruise-fans

Sequence of three computer programs predicts aerodynamic interference on lifting surfaces of transport-type aircraft which are equipped with lift and cruise fans; for example, high-bypass-ratio engine and wing-pylon tail configuration or fuselage-mounted lift-fan and wing-tail configuration

Calculation of the longitudinal aerodynamic characteristics of STOL aircraft with externally-blown jet-augmented flaps

A theoretical investigation was made to develop methods for predicting the longitudinal aerodynamic characteristics of externally-blown, jet-augmented wing-flap combinations. A potential flow analysis was used to develop two models: a wing-flap lifting surface model and a high-bypass-ratio turbofan engine wake model. Use of these two models in sequence provides for calculation of the wing-flap load distribution including the influence of the engine wake. The method can accommodate multiple engines per wing panel and part-span flaps but is limited to the case where the flow and geometry of the configuration are symmetric about a vertical plane containing the wing root chord. Comparisons of predicted and measured lift and pitching moment on unswept and swept wings with one and two engines per panel and with various flap deflection angles indicate satisfactory prediction of lift and moment for flap deflections up to 30 to 40 degrees. At higher flap angles with and without power, the method begins to overpredict lift, due probably to the appearance of flow separation on the flaps

Program LRCDM2: Improved aerodynamic prediction program for supersonic canard-tail missiles with axisymmetric bodies

Program LRCDM2 was developed for supersonic missiles with axisymmetric bodies and up to two finned sections. Predicted are pressure distributions and loads acting on a complete configuration including effects of body separated flow vorticity and fin-edge vortices. The computer program is based on supersonic panelling and line singularity methods coupled with vortex tracking theory. Effects of afterbody shed vorticity on the afterbody and tail-fin pressure distributions can be optionally treated by companion program BDYSHD. Preliminary versions of combined shock expansion/linear theory and Newtonian/linear theory have been implemented as optional pressure calculation methods to extend the Mach number and angle-of-attack ranges of applicability into the nonlinear supersonic flow regime. Comparisons between program results and experimental data are given for a triform tail-finned configuration and for a canard controlled configuration with a long afterbody for Mach numbers up to 2.5. Initial tests of the nonlinear/linear theory approaches show good agreement for pressures acting on a rectangular wing and a delta wing with attached shocks for Mach numbers up to 4.6 and angles of attack up to 20 degrees

Improvements to the missile aerodynamic prediction code DEMON3

The computer program DEMON3 was developed for the aerodynamic analysis of nonconventional supersonic configurations comprising a body with noncircular cross section and up to two wing or fin sections. Within a wing or fin section, the lifting surfaces may be cruciform, triform, planar, or low profile layouts; the planforms of the lifting surfaces allow for breaks in sweep. The body and fin sections are modeled by triplet and constant u-velocity panels, respectively, accounting for mutual body-fin interference. Fin thickness effects are included for the use of supersonic planar source panels. One of the unique features of DEMON3 is the modeling of high angle of attack vortical effects associated with the lifting surfaces and the body. In addition, shock expansion and Newtonian pressure calculation methods can be optionally engaged. These two dimensional nonlinear methods are augmented by aerodynamic interference determined from the linear panel methods. Depending on the geometric details of the body, the DEMON3 program can be used to analyze nonconventional configurations at angles of attack up to 25 degrees for Mach numbers from 1.1 to 6. Calculative results and comparisons with experimental data demonstrate the capabilities of DEMON3. Limitations and deficiencies are listed

Rolling moments in a trailing vortex flow field

Pressure distributions are presented which were measured on a wing in close proximity to a tip vortex of known structure generated by a larger, upstream semispan wing. Overall loads calculated by integration of these pressures are checked by independent measurements made with an identical model mounted on a force balance. Several conventional methods of wing analysis are used to predict the loads on the following wing. Strip theory is shown to give uniformly poor results for loading distribution, although predictions of overall lift and rolling moment are sometimes acceptable. Good results are obtained for overall coefficients and loading distribution by using linearized pressures in vortex-lattice theory in conjunction with a rectilinear vortex. The equivalent relation from reverse-flow theory that can be used to give economic predictions for overall loads is presented

Assessment of a wake vortex flight test program

A proposed flight test program to measure the characteristics of wake vortices behind a T-33 aircraft was investigated. A number of facets of the flight tests were examined to define the parameters to be measured, the anticipated vortex characteristics, the mutual interference between the probe aircraft and the wake, the response of certain instruments to be used in obtaining measurements, the effect of condensation on the wake vortices, and methods of data reduction. Recommendations made as a result of the investigation are presented

Computer programs for calculating pressure distributions including vortex effects on supersonic monoplane or cruciform wing-body-tail combinations with round or elliptical bodies

Computer programs are presented which are capable of calculating detailed aerodynamic loadings and pressure distributions acting on pitched and rolled supersonic missile configurations which utilize bodies of circular or elliptical cross sections. The applicable range of angle of attack is up to 20 deg, and the Mach number range is 1.3 to about 2.5. Effects of body and fin vortices are included in the methods, as well as arbitrary deflections of canard or fin panels

Experimental Stage Separation Tool Development in NASA Langley's Aerothermodynamics Laboratory

As part of the research effort at NASA in support of the stage separation and ascent aerothermodynamics research program, proximity testing of a generic bimese wing-body configuration was conducted in NASA Langley's Aerothermodynamics Laboratory in the 20-Inch Mach 6 Air Tunnel. The objective of this work is the development of experimental tools and testing methodologies to apply to hypersonic stage separation problems for future multi-stage launch vehicle systems. Aerodynamic force and moment proximity data were generated at a nominal Mach number of 6 over a small range of angles of attack. The generic bimese configuration was tested in a belly-to-belly and back-to-belly orientation at 86 relative proximity locations. Over 800 aerodynamic proximity data points were taken to serve as a database for code validation. Longitudinal aerodynamic data generated in this test program show very good agreement with viscous computational predictions. Thus a framework has been established to study separation problems in the hypersonic regime using coordinated experimental and computational tools