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NASA NTRSDate not supplied

A Methodology for Actuating RCS Jets in a Continuous, Time-Accurate CFD Simulation

Computational fluid dynamics simulations of Mars entry vehicle reaction control systems have been used to gain a better understanding of the wide range of flow phenomena encountered during various phases of flight. One aspect of this is simulating the system during controlled flight. Current practices model the system in a continuously actuated state, but this methodology is unable to accurately model the dynamic behavior of the system during EDL. One method to improve this is to create a coupled CFD-RBD-Control simulation. The first step to this is establishing a method of accurately actuating an RCS jet during a time accurate CFD simulation. This paper presents a method to achieve this actuation. There are three phases to the methodology: establishing the off state, creating a ramping function that will initialize the jet flow, and establishing the on state. A test case of a jet in a supersonic crossflow is implemented to establish the success of the methodology in being able to actuate a jet in the midst of an ongoing CFD simulation. It shows significant improvement in boundary condition responsiveness for linear and exponential ramping functions in comparison to the step function. A preliminary demonstration of the methodology is presented for the MSL vehicle which shows significant advantages for the implementation of a linear ramping function over a step function.

Alexandra Hickey↗

NASA NTRS1995

A numerical model for dynamic wave rotor analysis

A numerical model has been developed which can predict the dynamic (and steady state) performance of a wave rotor, given the geometry and time dependent boundary conditions. The one-dimensional, perfect gas, CFD based code tracks the gasdynamics in each of the wave rotor passages as they rotate past the various ducts. The model can operate both on and off-design, allowing dynamic behavior to be studied throughout the operating range of the wave rotor. The model accounts for several major loss mechanisms including finite passage opening time, fluid friction, heat transfer to and from the passage walls, and leakage to and from the passage ends. In addition, it can calculate the amount of work transferred to and from the fluid when the flow in the ducts is not aligned with the passages such as occurs in off-design operation. Since it is one-dimensional, the model runs reasonably fast on a typical workstation. This paper will describe the model and present the results of some transient calculations for a conceptual four port wave rotor designed as a topping cycle for a small gas turbine engine.

Paxson, D. E.↗

NASA NTRS1994

A fundamental aeroservoelastic study combining unsteady CFD with adaptive control

This paper describes a two-dimensional aeroservoelastic study in the time domain. The model, which is based on exact inviscid aerodynamics, correctly represents the large amplitude motions and the associated strong shock dynamics in the transonic regime. The aeroservoelastic system consists of a two degree-of-freedom airfoil with a trailing edge control surface. Using first-order actuator dynamics, a digital adaptive controller is applied to provide active flutter suppression. Comparisons between time-responses of the open-loop and closed loop systems show the ability of the trailing edge control surface to suppress non-linear transonic aeroelastic phenomena. A relation between actuator dynamics, sampling time-step and limits on the flap deflection angle to guarantee the effectiveness of the adaptive controller was demonstrated by the results generated.

Friedmann, P.↗

NASA NTRS1992

Pressure measurements on a rectangular wing with a NACA0012 airfoil during conventional flutter

The Structural Dynamics Division at NASA LaRC has started a wind tunnel activity referred to as the Benchmark Models Program. The primary objective of the program is to acquire measured dynamic instability and corresponding pressure data that will be useful for developing and evaluating aeroelastic type CFD codes currently in use or under development. The program is a multi-year activity that will involve testing of several different models to investigate various aeroelastic phenomena. The first model consisted of a rigid semispan wing having a rectangular planform and a NACA 0012 airfoil shape which was mounted on a flexible two degree-of-freedom mount system. Two wind-tunnel tests were conducted with the first model. Several dynamic instability boundaries were investigated such as a conventional flutter boundary, a transonic plunge instability region near Mach = 0.90, and stall flutter. In addition, wing surface unsteady pressure data were acquired along two model chords located at the 60 to 95-percent span stations during these instabilities. At this time, only the pressure data for the conventional flutter boundary is presented. The conventional flutter boundary and the wing surface unsteady pressure measurements obtained at the conventional flutter boundary test conditions in pressure coefficient form are presented. Wing surface steady pressure measurements obtained with the model mount system rigidized are also presented. These steady pressure data were acquired at essentially the same dynamic pressure at which conventional flutter had been encountered with the mount system flexible.

Rivera, Jose A., Jr.↗

NASA NTRS1992

NACA 0012 benchmark model experimental flutter results with unsteady pressure distributions

The Structural Dynamics Division at NASA Langley Research Center has started a wind tunnel activity referred to as the Benchmark Models Program. The primary objective of the program is to acquire measured dynamic instability and corresponding pressure data that will be useful for developing and evaluating aeroelastic type CFD codes currently in use or under development. The program is a multi-year activity that will involve testing of several different models to investigate various aeroelastic phenomena. This paper describes results obtained from a second wind tunnel test of the first model in the Benchmark Models Program. This first model consisted of a rigid semispan wing having a rectangular planform and a NACA 0012 airfoil shape which was mounted on a flexible two degree-of-freedom mount system. Experimental flutter boundaries and corresponding unsteady pressure distribution data acquired over two model chords located at the 60 and 95 percent span stations are presented.

Rivera, Jose A., Jr.↗

NASA NTRS2016

Improving Fidelity of Launch Vehicle Liftoff Acoustic Simulations

Launch vehicles experience high acoustic loads during ignition and liftoff affected by the interaction of rocket plume generated acoustic waves with launch pad structures. Application of highly parallelized Computational Fluid Dynamics (CFD) analysis tools optimized for application on the NAS computer systems such as the Loci/CHEM program now enable simulation of time-accurate, turbulent, multi-species plume formation and interaction with launch pad geometry and capture the generation of acoustic noise at the source regions in the plume shear layers and impingement regions. These CFD solvers are robust in capturing the acoustic fluctuations, but they are too dissipative to accurately resolve the propagation of the acoustic waves throughout the launch environment domain along the vehicle. A hybrid Computational Fluid Dynamics and Computational Aero-Acoustics (CFD/CAA) modeling framework has been developed to improve such liftoff acoustic environment predictions. The framework combines the existing highly-scalable NASA production CFD code, Loci/CHEM, with a high-order accurate discontinuous Galerkin (DG) solver, Loci/THRUST, developed in the same computational framework. Loci/THRUST employs a low dissipation, high-order, unstructured DG method to accurately propagate acoustic waves away from the source regions across large distances. The DG solver is currently capable of solving up to 4th order solutions for non-linear, conservative acoustic field propagation. Higher order boundary conditions are implemented to accurately model the reflection and refraction of acoustic waves on launch pad components. The DG solver accepts generalized unstructured meshes, enabling efficient application of common mesh generation tools for CHEM and THRUST simulations. The DG solution is coupled with the CFD solution at interface boundaries placed near the CFD acoustic source regions. Both simulations are executed simultaneously with coordinated boundary condition data exchange.

Liever, Peter↗

NASA NTRSDate not supplied

Numerical Investigation of Film Coefficient Engineering Methodology for Dissociated, Chemically Reacting Boundary Layers

Aerothermal analysis of spacecraft planetary entry is heavily dependent on heritage engineering models. The film coefficient heat transfer model examined in this paper estimates the convective heating to the vehicle for a laminar, dissociated, chemically reacting boundary layer for an Earth atmosphere. This model requires information about the vehicle and flow field for a given trajectory point and estimates a proportional relationship between enthalpy potential and convective heat flux. In practice it is the aerothermal engineer who must decide which assumptions are appropriate for their application. This work looks at numerous CFD simulations for an arbitrary, axisymmetric flight vehicle to analyze the relative importance of both the mass and energy constraints imposed at the wall boundary, as well as the effect of various diffusion models. The sensitivity of the engineering model is assessed based on the impact of the CFD assumptions as an effort to make it available to design engineers across industry.

numerical boundary layer↗

NASA NTRS1996

Spray Combustion Modeling with VOF and Finite-Rate Chemistry

A spray atomization and combustion model is developed based on the volume-of-fluid (VOF) transport equation with finite-rate chemistry model. The gas-liquid interface mass, momentum and energy conservation laws are modeled by continuum surface force mechanisms. A new solution method is developed such that the present VOF model can be applied for all-speed range flows. The objectives of the present study are: (1) to develop and verify the fractional volume-of-fluid (VOF) cell partitioning approach into a predictor-corrector algorithm to deal with multiphase (gas-liquid) free surface flow problems; (2) to implement the developed unified algorithm in a general purpose computational fluid dynamics (CFD) code, Finite Difference Navier-Stokes (FDNS), with droplet dynamics and finite-rate chemistry models; and (3) to demonstrate the effectiveness of the present approach by simulating benchmark problems of jet breakup/spray atomization and combustion. Modeling multiphase fluid flows poses a significant challenge because a required boundary must be applied to a transient, irregular surface that is discontinuous, and the flow regimes considered can range from incompressible to highspeed compressible flows. The flow-process modeling is further complicated by surface tension, interfacial heat and mass transfer, spray formation and turbulence, and their interactions. The major contribution of the present method is to combine the novel feature of the Volume of Fluid (VOF) method and the Eulerian/Lagrangian method into a unified algorithm for efficient noniterative, time-accurate calculations of multiphase free surface flows valid at all speeds. The proposed method reformulated the VOF equation to strongly couple two distinct phases (liquid and gas), and tracks droplets on a Lagrangian frame when spray model is required, using a unified predictor-corrector technique to account for the non-linear linkages through the convective contributions of VOF. The discontinuities within the sharp interface will be modeled as a volume force to avoid stiffness. Formations of droplets, tracking of droplet dynamics and modeling of the droplet breakup/evaporation, are handled through the same unified predictor-corrector procedure. Thus the new algorithm is non-iterative and is flexible for general geometries with arbitrarily complex topology in free surfaces. The FDNS finite-difference Navier-Stokes code is employed as the baseline of the current development. Benchmark test cases of shear coaxial LOX/H2 liquid jet with atomization/combustion and impinging jet test cases are investigated in the present work. Preliminary data comparisons show good qualitative agreement between data and the present analysis. It is indicative from these results that the present method has great potential to become a general engineering design analysis and diagnostics tool for problems involving spray combustion.

Chen, Yen-Sen↗

NASA NTRS2017

Free-Flight Trajectory Simulation of the ADEPT Sounding Rocket Test Using CFD

A computational study of the Adaptive Deployable Entry and Placement Technology (ADEPT) Sounding Rocket (SR-1) Test is presented using the US3D flow solver. ADEPT SR-1 is intended, in part, to assess the dynamic stability of this entry vehicle architecture. Given that no dynamic stability data exists for the ADEPT geometry, a limited ballistic range campaign has been performed to characterize the vehicle's stability characteristics pre-flight for Mach numbers between 1.21 and 2.5. Here, this data is used to assess the accuracy of US3D's free-flight CFD capability. Computed trajectories from US3D and experimental data show that the flow solver compares well in vehicle oscillation frequency, downrange distance, and oscillatory amplitude during high Mach number flight (Mavg = 2.36). For Mach numbers below 1.5, the solver under predicts total angle-of-attack by an average of 16%, but compares well in oscillatory frequency and downrange distance. Additionally, a capability for simulating the trajectory of the flight article through the atmosphere using CFD is presented. This capability couples US3D's free-flight capability to an atmosphere model that accounts for changes in free-stream density and temperature as the vehicle descends. Two simulations for the purpose of demonstrating the capability and viability of this approach are applied to SR-1 flight article, and some unique challenges are discussed.

CFD↗

NASA NTRS1983

A perspective of theoretical and applied computational fluid dynamics

The present effectiveness and future promise of computational fluid dynamics (CFD) are discussed. The CFD ideal of computer-designed flight vehicles and the progress that has been made toward its achievement, expressed in terms of primary and secondary pacing items, are also addressed. The primary pacing items listed by Chapman - grid generation, turbulence modeling, solution methodology development, and mainframe computer methods and architecture - are modified and reviewed. Important secondary pacing items in CFD development are also examined: algorithm development, complex geometry definition, and predata and postata processing. In addition, user requirements placed on CFD computer codes are defined and some problems associated with the dissemination of CFD technology are identified.

Kutler, P.↗

NASA NTRS1987

Current CFD efforts in projectile aerodynamics

Information is given in viewgraph form on current computational fluid dynamics (CFD) efforts in projectile aerodynamics. Topics covered include spinning projectiles, fin stabilized projectiles, model geometry, the variation of base drag with base bleed, the variation of normal force with Mach number, and chordwise pressure distribution.

Nietubicz, Charles J.↗

NASA NTRS1994

NASA geometry data exchange specification for computational fluid dynamics (NASA IGES)

This document specifies a subset of an existing product data exchange specification that is widely used in industry and government. The existing document is called the Initial Graphics Exchange Specification. This document, a subset of IGES, is intended for engineers analyzing product performance using tools such as computational fluid dynamics (CFD) software. This document specifies how to define mathematically and exchange the geometric model of an object. The geometry is represented utilizing nonuniform rational B-splines (NURBS) curves and surfaces. Only surface models are represented; no solid model representation is included. This specification does not include most of the other types of product information available in IGES (e.g., no material properties or surface finish properties) and does not provide all the specific file format details of IGES. The data exchange protocol specified in this document is fully conforming to the American National Standard (ANSI) IGES 5.2.

Blake, Matthew W.↗

NASA NTRS1995

Rapid Airplane Parametric Input Design (RAPID)

RAPID is a methodology and software system to define a class of airplane configurations and directly evaluate surface grids, volume grids, and grid sensitivity on and about the configurations. A distinguishing characteristic which separates RAPID from other airplane surface modellers is that the output grids and grid sensitivity are directly applicable in CFD analysis. A small set of design parameters and grid control parameters govern the process which is incorporated into interactive software for 'real time' visual analysis and into batch software for the application of optimization technology. The computed surface grids and volume grids are suitable for a wide range of Computational Fluid Dynamics (CFD) simulation. The general airplane configuration has wing, fuselage, horizontal tail, and vertical tail components. The double-delta wing and tail components are manifested by solving a fourth order partial differential equation (PDE) subject to Dirichlet and Neumann boundary conditions. The design parameters are incorporated into the boundary conditions and therefore govern the shapes of the surfaces. The PDE solution yields a smooth transition between boundaries. Surface grids suitable for CFD calculation are created by establishing an H-type topology about the configuration and incorporating grid spacing functions in the PDE equation for the lifting components and the fuselage definition equations. User specified grid parameters govern the location and degree of grid concentration. A two-block volume grid about a configuration is calculated using the Control Point Form (CPF) technique. The interactive software, which runs on Silicon Graphics IRIS workstations, allows design parameters to be continuously varied and the resulting surface grid to be observed in real time. The batch software computes both the surface and volume grids and also computes the sensitivity of the output grid with respect to the input design parameters by applying the precompiler tool ADIFOR to the grid generation program. The output of ADIFOR is a new source code containing the old code plus expressions for derivatives of specified dependent variables (grid coordinates) with respect to specified independent variables (design parameters). The RAPID methodology and software provide a means of rapidly defining numerical prototypes, grids, and grid sensitivity of a class of airplane configurations. This technology and software is highly useful for CFD research for preliminary design and optimization processes.

Smith, Robert E.↗

NASA NTRS2007

The Effect of Acoustic Disturbances on the Operation of the Space Shuttle Main Engine Fuel Flowmeter

The Space Shuttle Main Engine (SSME) uses a turbine fuel flowmeter (FFM) in its Low Pressure Fuel Duct (LPFD) to measure liquid hydrogen flowrates during engine operation. The flowmeter is required to provide accurate and robust measurements of flow rates ranging from 10000 to 18000 GPM in an environment contaminated by duct vibration and duct internal acoustic disturbances. Errors exceeding 0.5% can have a significant impact on engine operation and mission completion. The accuracy of each sensor is monitored during hot-fire engine tests on the ground. Flow meters which do not meet requirements are not flown. Among other parameters, the device is screened for a specific behavior in which a small shift in the flow rate reading is registered during a period in which the actual fuel flow as measured by a facility meter does not change. Such behavior has been observed over the years for specific builds of the FFM and must be avoided or limited in magnitude in flight. Various analyses of the recorded data have been made prior to this report in an effort to understand the cause of the phenomenon; however, no conclusive cause for the shift in the instrument behavior has been found. The present report proposes an explanation of the phenomenon based on interactions between acoustic pressure disturbances in the duct and the wakes produced by the FFM flow straightener. Physical insight into the effects of acoustic plane wave disturbances was obtained using a simple analytical model. Based on that model, a series of three-dimensional unsteady viscous flow computational fluid dynamics (CFD) simulations were performed using the MSFC PHANTOM turbomachinery code. The code was customized to allow the FFM rotor speed to change at every time step according to the instantaneous fluid forces on the rotor, that, in turn, are affected by acoustic plane pressure waves propagating through the device. The results of the simulations show the variation in the rotation rate of the flowmeter due to the interaction of the flow straightener wakes and the upstream propagating acoustic waves. A detailed analysis of the acoustic disturbance effects is presented along with an assessment of the impact on measurement accuracy.

Marcu, Bogdan↗

NASA NTRS2011

Compliant Turbomachine Sealing

Sealing interface materials and coatings are sacrificial, giving up their integrity for the benefit of the component. Seals that are compliant while still controlling leakage, dynamics, and coolant flows are sought to enhance turbomachine performance. Herein we investigate the leaf-seal configuration. While the leaf seal is classified as contacting, a ready modification using the leaf-housing arrangement in conjunction with an interface film rider (a bore seal, for example) provides for a film-riding noncontact seal. The leaf housing and leaf elements can be made from a variety of materials from plastic to ceramic. Four simplistic models are used to identify the physics essential to controlling leakage. Corroborated by CFD, these results provide design parameters for applications to within reasonable engineering certainty. Some potential improvements are proposed.

Hendricks, R. C.↗

NASA NTRS2012

Algebraic Turbulence-Chemistry Interaction Model

The results of a series of Perfectly Stirred Reactor (PSR) and Partially Stirred Reactor (PaSR) simulations are compared to each other over a wide range of operating conditions. It is found that the PaSR results can be simulated by a PSR solution with just an adjusted chemical reaction rate. A simple expression has been developed that gives the required change in reaction rate for a PSR solution to simulate the PaSR results. This expression is the basis of a simple turbulence-chemistry interaction model. The interaction model that has been developed is intended for use with simple one-step global reaction mechanisms and for steady-state flow simulations. Due to the simplicity of the model there is very little additional computational cost in adding it to existing CFD codes.

Norris, Andrew T.↗

NASA NTRS2016

RotCFD Analysis of the AH-56 Cheyenne Hub Drag

In 2016, the U.S. Army Aviation Development Directorate (ADD) conducted tests in the U.S. Army 7- by 10- Foot Wind Tunnel at NASA Ames Research Center of a nonrotating 2/5th-scale AH-56 rotor hub. The objective of the tests was to determine how removing the mechanical control gyro affected the drag. Data for the lift, drag, and pitching moment were recorded for the 4-bladed rotor hub in various hardware configurations, azimuth angles, and angles of attack. Numerical simulations of a selection of the configurations and orientations were then performed, and the results were compared with the test data. To generate the simulation results, the hardware configurations were modeled using Creo and Rhinoceros 5, three-dimensional surface modeling computer-aided design (CAD) programs. The CAD model was imported into Rotorcraft Computational Fluid Dynamics (RotCFD), a computational fluid dynamics (CFD) tool used for analyzing rotor flow fields. RotCFD simulation results were compared with the experimental results of three hardware configurations at two azimuth angles, two angles of attack, and with and without wind tunnel walls. The results help validate RotCFD as a tool for analyzing low-drag rotor hub designs for advanced high-speed rotorcraft concepts. Future work will involve simulating additional hub geometries to reduce drag or tailor to other desired performance levels.

AH-56 Cheyenne↗

NASA NTRS2016

Simulation of Sweep-Jet Flow Control, Single Jet and Full Vertical Tail

This work is a simulation technology demonstrator, of sweep jets used to suppress boundary layer separation and increase maximum achievable load coefficients. A sweep jet is a discrete Coanda jet that oscillates in the plane parallel to an aerodynamic surface. It injects mass and momentum in the approximate stream wise direction. It also generate turbulent eddies at the oscillation frequency, which are typically large relative to boundary layer turbulence, and which augmenting mixing across the boundary layer to attack flow separation. Simulations of a fluidic oscillator, the sweep jet emerging from the oscillator, and the suppression of boundary layer separation by an array of sweep jets are performed. Simulation results are compared to data from a dedicated CFD validation experiment of a single oscillator and its sweep jet, and from a study of a full-scale Boeing 757 vertical tail augmented with an array of sweep jets.2, 20 A critical step in the work is the development of realistic time-dependent sweep-jet in flow boundary conditions, derived from the results of the single-oscillator simulations, which create the sweep jets in the full-tail simulations. Simulations were performed using the Over flow CFD solver, with high-order spatial discretization and a range of turbulence modeling. Good results were obtained for all flows simulated, when suitable turbulence modeling was used.

flow control↗