NASA Technical Reports Server (NTRS) 20050205855: Towards Rocket Engine Components with Increased Strength and Robust Operating Characteristics

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by  NASA Technical Reports Server (NTRS)

AIAA 2005-4449
Towards Rocket Engine Components with
Increased Strength and Robust Operating
Characteristics
Bogdan Marcu
Ali Hadid
Pei Lin
Daniel Balcazar
The Boeing Company
Rocketdyne Propulsion and Power
6633 Canoga Avenue
Canoga Park, California
Man Mohan Rai
NASA
'
Ames Research Center
Moffett Field, California
Daniel J. Dorney
b
NASA
Marshall Space Flight Center
Huntsville, Alabama
41 AlAA/ASME/SAE/ASEE Joint Propulsion
st
Conference & Exhibit
10-13 July 2005
Tucson, Arizona
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Towards Rocket Engine Components with Increased Strength and Robust
Operating Characteristics
Bogdan Marcu, Ali Hadid, Pei Lin, Daniel Balcazar
The Boeing Company
Rocketdyne Propulsion and Power
6633 Canoga Avenue
Canoga Park, California
Man Mohan Rai
NASA
Ames Research Center
Moffett Field, California
Daniel J. Dorney
NASA
Marshall Space Flight Center
Huntsville, Alabama
Abstract WILD CA T/CORSAIR (NASA), FLUENT
(commercial), TIDAL (Boeing Rocketdyne) and, a
High-energy rotating machines, powering liquid
new family (AardvarWPhantom) of CFD analysis
propellant rocket engines, are subject to various
codes developed at NASA MSFC employing LOX
sources of high and low cycle fatigue generated
fluid properties and a Generalized Equation Set
by unsteady flow phenomena. Given the
formulation. Extensive aerodynamic performance
tremendous need for reliability in a sustainable
analysis and stress analysis carried out at Boeing
space exploration program, a fundamental
Rocketdyne and NASA MSFC indicate that the
change in the design methodology for engine
redesign objectives have been fully met. The paper
components is required for both launch and
presents the results of the assessment analysis
space based systems. A design optimization
and discusses the future potential of robust optimal
system based on neural-networks has been
design for rocket engine components.
applied and demonstrated in the redesign of the
Space Shuttle Main Engine (SSME) Low Copyrights etc Man lnserf 1.
Pressure Oxidizer Turbo Pump (LPOTP) turbine
1. Introduction
nozzle. One objective of the redesign effort was
to increase airfoil thickness and thus increase its One of the key requirements for sustainability of
strength while at the same time detuning the Space Exploration is the operational reliability and
vane natural frequency modes from the vortex robustness of the systems deployed and used in
shedding frequency. The second objective was space. In particular, the operation of propulsion
to reduce the vortex shedding amplitude. The systems based exclusively in space essentially
third objective was to maintain this low shedding requires systems which need minimal or no
amplitude even in the presence of large maintenance. This requirement forces the
manufacturing tolerances. All of these objectives elimination of the current philosophy of operation
were achieved without generating any for re-usable equipment, which is based on
detrimental effects on the downstream flow overhaul and repair at intervals in time dictated by
through the turbine, and without introducing any the limits in the reliability of the system’s
penalty in performance. The airfoil redesign and components. In space there is limited or no access
preliminary assessment was performed in the to repair or replace worn or damaged components.
Exploration Technology Directorate at NASA The absence of robust and reliable components will
ARC. Boeing/Rocketdyne and NASA MSFC necessitate propulsion systems based on high
independently performed final CFD redundancy with the associated cost and weight
assessments of the design. Four different CFD penalties.
codes were used in this process. They include
These shortcomings can be avoided by a associated with the trailing edge vortices are
change in design paradigm for the components insufficiently understood and consequently not yet
of space based propulsion systems. Here we properly modeled. Sondak and Dorney [7, 81
propose a major role for formal multi-objective showed that for a typical turbine cascade, the
optimization methods in design optimization by correct simulation of the flow separation at the
embedding them in the standard design trailing edge requires a grid of significant density
process. In order to obtain component designs (1 19 points at the trailing edge region). Also, in a
that are simultaneously characterized by high stage configuration, the shedding frequency may
performance, and high strength and robustness, either lock onto an upper harmonic of the blade
the designer must cope simultaneously with a passing fundamental frequency, or split into
large number of design variables. The human amplitudes at several harmonics.
mind can only simultaneously address a limited
While a large body of published literature
number of such variables in the context of a
addresses gas or air flow measurements, there is
conventional design process. Formal
little available material regarding vortex shedding
optimization algorithms can handle a large
on cascades of airfoilshydrofoils operating in liquid
number of design parameters and meet
flows. Lee, Hah and Loellbach [9] performed a
competing design requirements and objectives
RANS CFD analysis of the unsteady flow
The particular project described in the present interaction inside an axial flow pump stage. While
report exemplifies how such an approach works blade load comparisons with experiments are
and demonstrates the ability of the proposed presented, quantitative data pertaining to trailing
design paradigm in providing a revolutionary edge vortex shedding is not included in this
increase in component strength, reliability, investigation. Busby et al [lo] present numerical
performance and robustness. Such and experimental results for the same geometry
improvements will be routinely required by future with more details on the vortex shedding
space based systems. associated with the rotor blade row of the axial-flow
pump. Their study shows rotor trailing edge
shedding frequencies varying with radius (and
2. General description blade thickness) and locking into the vane passing
frequency harmonics. Ciocan et al [ll] published
High energy rotating machinery employed in
detailed optical measurements (LDV and PIV) of
liquid propellant rocket engines are subject to
the flow through a radial hydraulic turbine, however,
many sources of high cycle fatigue. Phenomena
without addressing the vortex shedding
such as cavitation, manifold flow instabilities,
phenomenon.
stator-rotor interaction, and vortex shedding
generate unsteady forces of various frequencies The present report addresses the issue of a very
subjecting the hardware to high unsteady stress energetic trailing edge vortex shedding
levels. phenomenon in an axial hydraulic turbine operating
in liquid oxygen, namely the turbine powering the
Particular attention has been focused on the
Low Pressure Oxidizer Turbo Pump (LPOTP) of the
phenomenon of vortex shedding at the trailing
Space Shuttle Main Engine (SSME). The analysis
edge of turbine airfoils. Most published studies
presented here has been made in the very
address cascade performance, with air or gas as
conservative context of addressing a flight safety
a working fluid. Han and Cox [l] present, in
issue for the Space Shuttle. In this study we
quantitative detail, the shedding phenomenon on
present the CFD calculations carried out in support
a turbine nozzle airfoil. They measure a range
of a revolutionary redesign of the LPOTP turbine
of Strouhal numbers at successively increasing
nozzle. Several CFD codes have been run
cascade discharge velocities. More and Adhye
simultaneously in order to determine and compare
[2] , Sieverding[3, 41 and Browand [5] showed
the frequency and amplitudes of shedding induced
that the shedding frequency and the size of the
pressure fluctuations for the nominal and
vortices are influenced by the nature of the
redesigned airfoils.
boundary layer. These vortices have a
significant effect on the shape and depth of the
wake. Furthermore, the characteristics of vortex
Insert 2 Man
shedding have an impact on the interaction with
the blade downstream. Contini et al [6] and
Sieverding [4] suggest that the unsteady effects
Based on the CFD calculations, structural fluxes are discretized according to the scheme
dynamics and stress calculations have been developed by Roe [12]. The viscous fluxes are
made in order to determine the expected safety calculated using standard central differences. An
factors and the expected life for the new approximate-factorization technique is used to
component,and the impact on the entire reduce and simplify the matrices which need to be
turbopump.. solved at every time step. Newton sub-iterations
are used at each time step to enhance stability and
Figure 1 shows a side view of the Space Shuttle
reduce linearization errors. The equations of motion
Main Engine (SSME), and the location of the
are extended to turbulent flows using an eddy
Low Pressure Oxidizer Turbo Pump (LPOTP).
viscosity formulation. The turbulent viscosity is
Figure 2 shows a simplified schematic of the calculated using the two-layer Baldwin-Lomax
engine. A cross section of the LPOTP is shown algebraic turbulence model [13]. The computational
in Fig. 3. The pump side of the LPOTP consists procedure uses 0- and H-type zonal grids to
of a single piece inducer powered by a six stage discretize the flow field and facilitate relative motion
hydraulic turbine. The turbine is fed via a tap-off between rotor and stator rows. The O-grids are
line from the discharge of the High Pressure body-fitted to the surfaces of the airfoils and
Oxidizer Turbo Pump (HPOTP). The discharge generated using an elliptic grid generator. They are
of the LPOTP turbine is re-circulated back to the used to accurately resolve the viscous flow in the
inlet of the HPOTP. blade passages and to easily apply the algebraic
turbulence model. The algebraically-generated H-
Figure 4 shows the exact location of the LPOTP
grids are used to discretize the remainder of the
turbine nozzle. Metallurgical inspections of the
flow field. Details of the algorithms and the gridding
nozzle parts have found evidence of high cycle
methodology can be found in references [14, 15
fatigue (HCF) at the nozzle trailing edge near
and IS].
the end walls. Analysis of the known sources of
excitation of HCF points to vortex shedding as Calls to Message Passing Interface (MPI) and
the most probable cause. Strong vortex OpenMP parallel computation libraries have been
shedding can generate flapping of the trailing implemented in the code to reduce the computation
edge as shown in the detail in Fig.4. Indeed, time for large-scale three-dimensional simulations.
CFD analyses which will be presented in the The use of MPI allows the coupling of different
next sections show that shedding frequencies geometric components, such as a turbine cavity, in
are close to the the blade trailing edge flapping a straightforward manner.
mode natural frequency..
The Corsair family of codes has been exhaustively
The LPOTP turbine nozzle component is validated for gas turbine analysis. In the present
currently being replaced at carefully monitored study however, the codes have been used to
time intervals ensuring full safety for the Shuttle analyze a turbine geometry operating in liquid.
flights. If the same turbopump were to operate While the input parameters have been carefully
within a space based system, the HCF wear of crafted to best approximate the actual flow
the part would limit the operational life of the conditions, the speed of sound in the working fluid
overall system. Clearly, it is imperative that could not be accurately scaled. Due to this
reliability and robustness are built into the limitation, analyses of the LPOTP turbine have
components for space based propulsion initially been limited to the first turbine stage; further
systems. expansion of the computational domain to include
more stages downstream would have led to density
3. CFD Analvsis
effects that would result in a departure of the
The flow analyses presented in this study have computed solution from the actual physics.
been performed using several CFD codes.The
With the above limitations in mind, a newer code
NASA MSFC Corsair family of codes were used
has been used for the present analysis. The NASA
during the initial phase of the project These
MSFC Aardvark code employs a
codes solve the time-dependent, three-
compressible/incompressible formulation based on
dimensional Reynolds-averaged Navier-Stokes
the Generalized Equation Set [17, 181 formulation
equations. The numerical algorithm used in the
for the Reynolds Averaged Navier Stokes
computational procedure consists of a time-
equations. A preconditioning algorithm is used for
marching, implicit, finite-difference scheme. The
incompressible flows. The new algorithm utilizes
procedure is spacially third-order accurate and
the true thermodynamic properties of the working
temporally second-order accurate. The inviscid
fluid. A library of real fluid properties has been second-order time stepping were used to reduce
implemented into the code. spatial and temporal discretization errors.
The results obtained with Corsair and Aardvark A refined unstructured mesh near the blade surface
are very similar, displaying the same trends and and downstream of the trailing edge together with a
changes in pressure fluctuation frequencies and small time step size were used in order to
amplitude. However there are small quantitative accurately capture the shedding frequency and
differences in surface pressures, shedding amplitude of the LPOTP turbine nozzle. Great care
frequencies and amplitudes. The two codes are was exercised to ensure grid and time-step
to some extent numerically related to each other independent solution by refining the grid and
(the airfoil topology is modeled with the same choosing a small time step of the order of 1/50 of a
type of grids, while the zonal boundary condition shedding period. The number of sub-iterations per
implementation follows similar logic, and the time step was chosen large enough to ensure
same turbulence model is used). Hence two convergence at each time step.
additional codes utilizing different turbulence
Turbulence effects were captured using the
models and based on unstructured grids have
unsteady Reynolds Averaged Navier-Stokes
also been used for analysis.
(URANS) full Reynolds stress model (RSM) of
The Tidal code, which stands for Time Iterative Launder, Reece and Rodi [19]. In this model the
Density/pressure based Algorithm, is a RANS full transport equations for the turbulent Reynolds
code developed at Rocketdyne. The code stresses are solved together with the continuity and
utilizes a finite volume, multi-zone method, and momentum equations. The sliding mesh capability
a steadyhnsteady modularized flow simulation of FLUENT is used to predict the time dependent
algorithm. A unified approach is employed to flow through a 2D rotor-stator blade row. The time-
combine the density- and pressure-based varying rotor-stator interaction is modeled by
methods to enable the computation of flow allowing the mesh associated with the moving rotor
fields ranging from incompressible to supersonic to translate (slide) relative to the stationary mesh
flows. In the present analysis, a dual time- associated with the stator blade. Initially a steady
stepping method is used to obtain a time- flow calculation with a stationary rotor was initiated
accurate solution. A central difference scheme and the solution obtained was used as a starting
is applied to the convection terms and viscous solution for the time dependent sliding-mesh
term. An adaptive second order dissipation calculation. Unsteady lift forces on the rotor and
method is employed for smoothing. The 2- the stator blades were monitored to determine
equation k-E turbulence model is used. when the unsteady flow predictions became time-
Precondition is used in the analysis. Parallel periodic and independent of the initial condition.
processing on a Linux cluster enables a rapid
4. Nominal Geometrv Analvsis. HCF
turn-around.
induced bv vortex sheddinq
The commercial code FLUENT has also been
Figure 5 shows the flow field through the LPOTP
used here for 2D single airfoil and 2D full stage
nozzle cascade obtained using Wildcat. Entropy is
analysis. FLUENT is a general purpose
used to visualize the vortex shedding patterns at
computer program for modeling fluid flow and
the airfoil’s trailing edge. The result was obtained
heat transfer in complex geometries. It provides
from a simulation of the flow through a 43 airfoil
complete mesh flexibility in solving flow
cascade, at a pressure drop corresponding to
problems with unstructured meshes in 2D and
109% Rated Power Level (RPL) of the SSME.
3D geometries using triangularhetrahedral,
Unlike cylinder vortex shedding, the shedding at the
quadrilateralhexagonal, or mixed (hybrid) grids
trailing edge of the nozzle creates alternating
that includes prisms and pyramids. The mesh
vortices of unequal strength: the vortex released
can be generated about complex geometries
from the pressure side of the vane is stronger. The
with relative ease using the preprocessor
effect of the pressure unsteadiness is felt upstream
package GAMBIT. FLUENT allows for multiple
of the trailing edges, as seen on the vane pressure
moving reference frames, including sliding mesh
loading envelopes in Fig. 6. Significant
interfaces and mixing planes for the modeling of
unsteadiness at the trailing edge region can be
rotor/stator interaction. The segregated solver
deduced from the large excursions in the pressure
formulation was used to solve the continuity,
envelope. The effect is felt on the suction side for
momentum and scalar equations sequentially.
almost two thirds of the blade chord from the
Second order spatial accuracy and implicit
trailing edge. The frequency and amplitude of the
I :
vortex shedding are shown on the FFT diagram with the two different codes are very close. The
in Fig. 7. Fifty points out of the 301 grid points Strouhal number based on local velocity at the
around the blade were sampled, and the trailing edge and the trailing edge thickness is St =
location with the maximum amplitude of 0.223-0.235. The pressure amplitude averaged
pressure fluctuation was selected (the point is over the 0.100 inch section at the trailing edge is
located at the tangency point where the trailing *72 psi (Aardvark simulation).
edge circle meets the pressure side of the
Figure 10 shows the nozzle vane response as a
airfoil). The results indicate a point pressure
function of frequency for the trailing edge flap
fluctuation of *181 psi at a frequency of 44,175
mode, and the spectral content of the vortex
Hz. This large amplitude diminishes rapidly as
shedding phenomenon at 109% engine Rated
one moves away from the point of maximum
Power Level (RPL). The vane peak resonance
amplitude (the corresponding tangency point on
response value displayed corresponds to the
the suction surface has a slightly lower
natural frequency of the nominal geometry as
amplitude at the same frequency). Hence a
measured in the lab at 47 kHz without corrections
more meaningful measure of unsteady stress is
for LOX mass flow effects which can account for
the frequency and amplitude of fluctuating
20%-40% reduction in frequency [21]. Additional
pressure averaged over a 0.100 inch width
uncertainty comes from the actual hardware which
section along the airfoil’s trailing edge. This
* has large deviations from the nominal airfoil
averaged measure has an amplitude of 57 psi
geometry. Overall, with all the corrections and
at the same frequency of 44,174 Hz.
variations due to hardware geometry deviations,
Figure 8 shows the flow field obtained from a 2- the blade natural frequency associated with the
D CFD Wildcat simulation of a stage trailing edge flap mode can range anywhere .
configuration. The first stage of the LPOTP’ between 24 kHz and 46 kHz. The vortex shedding
turbine is modeled with 44 nozzles and 66 frequency (43.6 kHz) in Figure 10 represents the
blades (2-nozzles and 3-blades) operating at value obtained with Aardvark for the nominal
109% RPL of the SSME. The vortex patterns geometry. Several CFD calculations have been
due to shedding are clearly visible in the wakes performed for nozzles with variations in geometry
all the way through the rotor row inlet. There, the representative of hardware variations in vane
vortices interact with the rotor blade, impinging trailing edge thicknesses. Furthermore, CFD
on the leading edges, and then being convected calculations for cylinders have been performed [20]
downstream through the rotor passage. The at similar Reynolds numbers (based on local TE
spectral content of the pressure fluctuation at velocity and TE diameter) in order to calibrate the
the tangency point on the pressure side of the code output for fluctuation frequencies. The results
nozzle is shown in Figure. 9. The results indicate an over-prediction of shedding frequency
obtained from the Wildcat and Aardvark by 1520%. These uncertainties in the computed
simulations are compared in this figure. The data and the hardware geometry result in a
data show fluctuating pressure amplitudes of shedding frequency range between 28 kHz and 45
&12 psi at a frequency of 45487 Hz from kHz. Thus the shedding frequency range and the
Wildcat and &08 psi at 43592 Hz from vane natural frequency range overlap. This overlap
Aardvark. The frequency and amplitude values and the lock-in mechanisms observed many times
are very close for the two codes. in operation [7, 8, 101 are a strong indication that
vortex shedding is the major cause for HCF wear
A comparison of the results obtained using
and damage observed in the operation of the
Wildcat for the cascade and stage configurations
LPOTP.
(Figure 7 and 9) show differences in shedding
characteristics. The shedding frequency for the
stage configuration is slightly different than the
5. The traditional redesiqn approach and
shedding frequency for a vane in isolation
its shortcominqs
because of a lock-in effect. The shedding
frequency locks on to the closest upper The initial approach taken to address the
harmonic of the vane passing fundamental resonance problem consisted of a “retrofit” of the
frequency, or to the closest upper harmonic and existing shelf parts. The details are discussed at
a half [lo]. In this case the Wildcat simulation length in reference [21]. At the expense of a small
indicates a lock into the 12thh armonic, while the percentage of the overall turbine performance, a
Aardvark simulation indicates a lock into the 1 lth shortening of the LPOTP nozzle airfoil via
+ It should be noted that the results obtained machining was considered for the purpose of
Y2.
decoupling the vortex shedding from the flap It is expected that the redesign of the baseline
mode of the vane TE. Shortening of the blade, airfoil using traditional design techniques will yield a
results in a thicker trailing edge, and new and thicker airfoil that has a lower shedding
consequently, the vortex shedding frequency frequency but a larger shedding amplitude. This
decreases, while the stiffness and hence the amplitude may not be as large as that produced by
trailing edge flap natural frequency increase. the cutback vane, but large enough to surface as a
disturbance amplitude at an undesired frequency
In the following we will briefly present the on the rotor blade downstream.
shortcomings of this retrofit approach.
13 insert Man
After several trade-off analyses, a vane cut of
0.100 inch in length at the trailing edge was
found to provide a sufficient increase in trailing 6. New, extended desiqn requirements
edge thickness. Figure 11 shows the baseline
The design requirements for a new nozzle airfoil
and retrofit first stage turbine geometries. The
should therefore contain additional provisions for
cut cannot be accomplished uniformly over the
preventing or limiting the introduction of unintended
entire vane span because the cutting tool has no
downstream perturbations. Obviously the overall
access near the end walls. Consequently there
turbine performance must be preserved in order to
is a round cut towards the walls.
ensure the proper operation of the turbopump
Figure 12 shows a snapshot of the flow field component.The complete set of design
through the turbine stage for both the baseline r.e quirements can be summarized as follows:
and retrofit geometries at midspan. The thicker
The airfoil should be as thick as possible at the
trailing edge of the retrofit nozzle and the sharp
TE, thereby increasing the natural flap mode
edges left after the cut (only limited chamfer can
frequency, and decreasing the shedding
be achieved) produce a lower vortex shedding
frequency, thus decoupling the two frequencies in
frequency (evidenced by the lower density of
.o peration
blobs convected downstream) but stronger
vortices (evidenced by the sharper color contrast Steady state (time averaged) flow conditions at
in the plot ###). nozzle row discharge should be preserved:
Figure 13 displays the quantitative effects of the
-same nozzle throat area in order to conserve
modification. On the one hand, the objective of
LOX mass flow value.
reducing the shedding frequency is achieved for
the nozzle vane, as shown in Figure 13a: the - same discharge flow angle in order to conserve
shedding frequency has been lowered by about . downstream-rotor blade work.
17 kHz, from 44 kHz to 28 kHz. Modal analysis
Additional flow disturbances should not be
and lab tests for the modified nozzle vane
introduced downstream:
indicate trailing edge flap mode natural
frequencies as high as 65-70 kHz, which - control of flow separation at the nozzle TE
reduces to about 52 kHz in operation after
- diminished the shedding amplitude
accounting for the LOX mass flow effect. Thus a
.
frequency separation of 24 kHz is achieved.
Airfoil shedding should be insensitive to geometry
Paragraph required. (couple to ###) discrepancies due to manufacturing or wear and
tear
Unfortunately, Figure 13b shows that the nozzle
-
problem has been solved at the expense of the Expected manufacturing method: casting with
rotor blade downstream. Stronger vortices shed f 0.006 inch tolerance with large variations
by the cut back vane at lower frequency are
The last requirement has been added as a result of
convected into the rotor passage. These vortices
large variations in the current hardware geometry
impinge on the rotor with enough energy to
with the intention of partially removing the
create a strong disturbance at a frequency of 29
operational performance sensitivity to airfoil
kHz. The frequency of this disturbance lies in
geometry deviations.
the middle of the frequency range for the rotor
blade first bending mode natural frequency. This This set of requirements is extremely complex, and
is not acceptable. Therefore, the retrofit solution difficult, if not impossible to achieve within the
has been rejected. context of traditional design techniques. Thus, it
constitutes an excellent case for formal design around the airfoil, starting and ending at the leading
optimization methodology. Drawing on previous edge, and the z axis represents the amplitude of
results from collaborations between NASA ARC the pressure fluctuation at each point. The plot
Exploration Technology Directorate and shows a reduction in shedding frequency from 43.1
Rocketdyne, the SSME LPOTP Nozzle redesign kHz to 32.7 kHz, and a remarkable reduction in the
team decided to address the redesign task by overall amount of unsteadiness for the 05 profile.
utilizing design optimization technology The new reduction provides a reduction of 76% in
developed at NASA ARC. The airfoil redesign peak amplitudes near the trailing edge.
and preliminary assessment was performed in
The data in figures 15 and 16 were obtained from
the Exploration Technology Directorate at NASA
Aardvark
ARC. Boeing/Rocketdyne and NASA MSFC
independently performed final CFD Figure 17 shows a similar comparison but obtained
assessments of the design from a 3-D Tidal simulation at midspan. In this
computation pressure fluctuations were recorded at
a location near the vane trailing edge, where
7. ODtimization methodoloqv. The new Aardvark and Wildcat showed maximum
airfoil geometry amplitudes for the baseline and 05 airfoils. The
plot shows the fluctuations for 05 are about 50%
[Man, insert 4 +figs]...
less than the fluctuations for the baseline, with a
Figure 14 shows the new design overlaid on the corresponding frequency decrease from 35 kHz to
baseline nozzle design for comparison. This 25 kHz.
geometry is the result of five design iterations
Figure 18 shows results obtained from 2-D
and is referred to as 05 in the rest of the text.
FLUENT computations, where data has been
The airfoil is very thick, and in particular, the
sampled as in the Tidal computation, Le. at a
trailing edge thickness is increased by about
location near the vane trailing edge, where
60% Intuitively one would expect such an
Aardvark and Wildcat showed maximum
increase in thickness to introduce a significant
amplitudes for the baseline and 05 airfoils. This
change in the wake profile and consequently a
plot shows a 52% reduction in amplitude for the 05
change in the flow downstream of the nozzle
vane with a corresponding reduction in frequency
row. A significant amount of analysis, presented
from 56 kHz to 47 kHz. FLUENT yields shedding
in the following sections, was performed to
frequencies which are clearly higher than physical,
determine if the new airfoil produced any
given the calibration runs made for cylinders using
detrimental changes in the flow downstream.
Aardvark (no calibration runs were made for
cylinders using TIDAL or FLUENT).
8. CFD Analvsis of performance The important feature in figures 16 through 18
is a consistent trend obtained with a variety of
8.1 Nozzle Performance
CFD codes. The 05 vane reduces pressure
Figure 15 presents a comparison of the nozzle fluctuation amplitudes by 50-75% while
vane pressure loading, for the baseline and reducing the shedding frequency by about 10
optimized airfoils. One observes that while the kHz.
baseline airfoil loading occurs mostly on the last
two thirds of the axial chord, the optimized airfoil
distributes the load more uniformly. A vertical 8.2 Overall Turbine Performance
line marks the proximity of the trailing edge in
Figures 19 and 20 compare the 1'' stage rotor
both plots. A significant difference exist in the
blade loading and 2"ds tage stator loading at 109%
pressure difference between the pressure side
engine RPL. Each plot overlays the pressure
and the suction side at the marked location; a
loading for the configurations using the baseline
larger Ap for the baseline geometry and a
and the 05 nozzles. Both plots indicate
diminished Ap for 05.
insignificant changes in rotor blade and stator vane
loadings downstream. It is reasonable to assume
Figure 16 is a carpet plot comparing the level of
that if only small changes can be observed in the
unsteadiness in pressure oscillations at different
locations on the airfoils. The plot should be two downstream rows following the replaced
nozzle, even smaller changes in performance are
interpreted as follows: the x axis represents
to be expected further downstream and therefore
frequency, in Hz, the y axis represents locations
the nozzle replacement produces little or no nozzle row when the current baseline nozzle
change in turbine performance geometry is replaced by the 05 airfoil. These
results have built confidence in the capabilities of
the new design.
8.3 Downstream Unsteadv Flow Analvsis
8.4 Robustness in operation
Figure 21 shows a comparison of the unsteady
(Man insert 5)
tangential force on the 1'' stage rotor blade
obtained from Aardvark. The FFT analysis for
the time varying tangential force is shown in this
9. Structural dvnamics analvsis.
figure. The fundamental frequency for the stage
configuration is the vane passing frequency of An exhaustive amount of analysis for assessing the
3.79 kHz corresponding to 44 nozzle vanes in structural dynamics of the new design has been
the simulation. The use of 05 instead of performed. Only the most important aspects are
baseline nozzle airfoil decreases the amplitude included in this paper.
at this frequency from *8.8% of the mean value
Figure 25 shows the TE flap mode response,
to *6.8%. Additionally, 05 does not introduce
corrected for a 30% LOX mass flow effect, for both
additional disturbances other than the small
the baseline and the 05 airfoils. The frequency
amplitude disturbances similar to those
ranges associated with the vortex shedding at
observed for the baseline design. This is a
109% engine RPL for the baseline and 05 airfoils
significant accomplishment given that the TE
are also shown in the figure. Flow data was
thickness of 05 is significantly larger that that of
obtained from Aardvark for airfoils at the extremes
the baseline airfoil.
of hardware geometry variation and adjusted by a
Given the importance of limiting or perhaps even 20% frequency shift based on calibration using
reducing the disturbances introduced cylinder simulations. The 05 performance is treated
downstream compared to the baseline case, the in a conservative manner, in the sense that
Tidal and FLUENT codes were also used in this although this geometry is characterized by
analysis. The results of these analyses are given exceptional robustness, and minimum variation in
below: flow performance due to geometry variation, the
same percent statistical bounds are applied on
Figure 22 shows the FFT analysis of the
vortex shedding frequencies as for the baseline
unsteady tangential force on the rotor blade
configuration. The plot clearly shows the good
obtained from Wildcat. The plot indicates a
separation in shedding and flap mode natural
reduction in amplitude from *19% to ~12%a t
frequencies for the 05 geometry at 109% engine
fundamental frequency and diminished
RPL regime. A comparison of figures 10 and 25
amplitudes at higher frequencies with the use of
shows the significant increase in frequency
05.
separation with the use of 05.
Figure 23 shows the same analysis using data
Figure 26 is an elaboration of the information in
produced by TIDAL. The amplitude at the
Figure 25 throughout the throttling range of the
fundamental frequency is practically unchanged
SSME engine. The plot shows the range of
in this calculation (a slight reduction from *4.2%
frequencies associated with the baseline geometry
to 23.8%) while the high frequency amplitudes
in blue fields and the ones associated with the 05
remain very small (less than 0.5% of average)
geometry in purple. CFD simulations have been
with the use of 05.
performed for the flow regimes corresponding to
In Figure 24, the analysis is repeated for data 64.5%, 8O%, 104% and 109% engine RPL, with
produced FLUENT. FLUENT shows no change intermediate RPL point performance obtained via a
in the amplitude at the fundamental frequency cubic spline fit. The data processed for the chart
(*19.5%) and no additional disturbances contains all the corrections for the uncertainties
introduced at higher frequencies with the use of related to the flow conditions, LOX mass flow
05. effects and possible alterations in the manufactured
geometry of the hardware, and thus the large range
The simulation results provided by codes with
of frequency values for both geometries. Again, the
different numerical schemes, different grid
data is processed taking the conservative approach
structures and different turbulence models all
typical for analysis of man-rated propulsion
display a consistent trend: there is a modest
hardware such as SSME. One can observe that
benefit or no change downstream of the
the baseline design geometry shedding The methodology presented here can be applied to
frequency ranges interfere significantly with the airfoil design for all turbomachinery equipment
natural vane frequency associated with the TE developed for space operations. Based on detailed
flap mode at all engine RPL regimes. In spite of analysis identifying all sources of intense stress,
the conservative estimate of the performance for LCF and HCF, including provisions necessary for
05, a large margin of 27% is obtained as deep throttling of rocket engines, design
separation between frequencies at the highest requirements similar to the extended design
RPL regime considered (109%). With this requirements presented in section 6 can be
margin, the frequencies are considered formulated. Subsequent airfoil design using
completely detuned. optimization methodology as presented here will
ensure increased strength and robust operating
characteristics for the components designed.
10. Stress analysis.
Given the powerful algorithms which form the
Similar to the structural dynamics section, only foundation of the optimization methodology
some principal results are presented here from demonstrated here, one can extend its application
the large amount of stress analysis done for this to other components of space based systems and
project. Besides pressure loads induced by the subsystems. Principal candidates are those
steady state and unsteady aerodynamic components deemed the most fragile in operation,
phenomena, additional loads generated by the for example, seals of various kinds, bearings,
engine vibration, static loads transmitted through injectors, or valves.
the turbopump support structure and mechanical
The same methodology can be applied on a larger
and thermal loads during transients have been
scale, and earlier in the evolution of a mission
included in the analysis.
conceptual design, for trade-off studies. Mission
Table 1 summarizes some of the main results formulation, vehicle/platform architecture and space
for the nozzle vane, showing the overall change flight procedures can be parameterized using large
in stress resulting from the replacement of the sets of variables modeling all necessary aspects.
baseline geometry with the 05 geometry. The Multi-objective optimization can be performed in
various categories of stress considered are high-dimensional design spaces that include all the
.
reduced by 19% to 600% generating an overall necessary variables “Optimize-not-compromise
”
increase in the safety factor from 3.5 to 6.3 and should become the motto of conceptual thinking for
thus giving the part fitted with the 05 airfoil an all future space exploration architectures.
essentially infinite life in operation.
Figure 27 shows the locations of peak
+ 12. Conclusions.
cumulative stress (steady state alternate) on
the rotor blade. The locations are near the A design methodology based on formal multi-
leading edge (Figure 27 a) and trailing edge objective optimization technology has been applied
(Figure 27 b). The maximum occurs in the and demonstrated for the redesign of the SSME
leading edge region where the calculated factor LPOTP turbine nozzle. The new nozzle design has
of safety is 4.91 for the baseline design achieved about 100% increase in strength,
configuration and 6.31 for the 05 configuration, significantly extended life in operation and an
an increase by 28% in safety. elevated robustness in operation while the overall
turbine performance has been maintained the same
These results constitute an essential element of
as for original design.
support for the design methodology for
components of space based propulsion The design methodology is proposed as a standard
systems. Without any penalty in performance, a design procedure for components of space based
component designed via multi-objective systems as it provides a means to design such
optimization methodology has resulted in almost components with significantly improved strength,
a 100% increase in strength, robust operation reliability and robust operating characteristics.
and practically infinite life.
13. References
11. Contribution potential for Space
Exlploration Missions 1 Han, L. S. and Cox, W. R., “A visual Study of
Turbine Blade Pressure-Side Boundary Layers”,