zotero-db/storage/R24JU7TV/.zotero-ft-cache

NASA Technical Memorandum 104330

Predicted Performance of a ThrustEnhanced SR-71 Aircraft with an
External Payload

Timothy R. Conners NASA Dryden Flight Research Edwards, Cafifornia

Center

1997

National Aeronautics and Space Administration
Dryden Flight Research Center Edwards, California 93523-0273

PREDICTED PERFORMANCE OF A THRUST-ENHANCED SR-71 AIRCRAFT WITH AN EXTERNAL PAYLOAD

Timothy R. Conners* NASA Dryden Flight Research
Edwards, California

Center

ABSTRACT
NASA Dryden Flight Research Center has completed a preliminary performance analysis of the SR-71 aircraft for use as a launch platform for high-speed research vehicles and for carrying captive experimental packages to high altitude and Mach number conditions. Externally mounted research platforms can significantly increase drag, limiting test time and, in extreme cases, prohibiting penetration through the high-drag, transonic flight regime.
To provide supplemental SR-71 acceleration, methods have been developed that could increase the thrust of the J58 turbojet engines. These methods include temperature and speed increases and augmentor nitrous oxide injection. The thrust-enhanced engines would allow the SR-71 aircraft to carry higher drag research platforms than it could without enhancement.
This paper presents predicted SR-71 performance with and without enhanced engines. A modified climb-dive technique is shown to reduce fuel consumption when flying through the transonic flight regime with a large external payload. Estimates are included of the maximum platform drag profiles with which the aircraft could still complete a high-speed research mission. In this case, enhancement was found to increase the SR-71 payload drag capability by 25 percent. The thrust enhancement techniques and performance prediction methodology are described.

NOMENCLATURE

DAFICS

digital automatic flight and inlet control

DPS

digital performance simulation

GHRV KEAS N20

generic hypersonic research
equivalent airspeed, knots nitrous oxide

vehicle

77"5

turbine exhaust temperature, deg F

WFAB

augmentor fuel flow, Ib/hr

system

*Aerospace Engineer.

INTRODUCTION There are several high-supersonic experiment packages current-
ly in design that call for flight test verification. These packages include advanced ramjet and rocket engine, inlet, and aerodynamic concepts for hypersonic vehicles, I as well as concepts that could benefit the high-speed commercial aviation research effort. The NASA Dryden Flight Research Center SR-71 aircraft (fig. l ) offer the capability to acquire long-duration flight test time at high altitudes and supersonic speeds in a captive-carry test environment. In cases where hypersonic Mach number data are required, the SR-7 l aircraft can feasibly provide air-launch capability for large-scale hypersonic research vehicles (fig. 2).
A structurally modified SR-71 aircraft can carry external payloads weighing up to 20,000 lbm (9072 kg). This large weight limit permits flexibility in the configuration of a research package. However, within this weight limit, it is easy to design an external payload package whose additional drag exceeds the excess thrust capability of an SR-71 aircraft using unmodified J58 engines. This effect was seen during feasibility studies performed by Lockheed Advanced Development Company (Palmdale, California) and NASA for a generic hypersonic research vehicle (GHRV). The GHRV was to be air-launched from the SR-71 vehicle and accelerate under its own power to speeds greater than Mach 10. The vehicle was to be used primarily for investigating scramjet engine performance.
Performance simulations had indicated that without additional excess thrust, the SR-71/GHRV combination could not accelerate through the transonic drag rise. Therefore, in support of the GHRV effort, Pratt & Whitney (West Palm Beach, Florida) performed a feasibility study under subcontract to Lockheed to increase the thrust of the J5g engines. 2 The goal of the study was a l0 percent net thrust increase within the transonic Mach range. This increase was deemed capable of providing adequate margin to allow the vehicle to reach its cruise condition. The study concluded that the goal was achievable by using compressor rotor speed and turbine

EC91-520-7
FIGURE 1. ONE OF THE THREE NASA DRYDEN SR-71 AIRCRAFT.

FIGURE 2. MODEL OF THE SR-71 AIRCRAFT VEHICLE.

AND BACK-MOUNTED

GENERIC HYPERSONIC

EC92-7241-3 RESEARCH

exhaust temperature, TT5, increases along with nitrous oxide, N20, augmentor injection.
The GHRV program did not proceed beyond the study phase. However, the feasibility of the engine modifications has made the enhancement technique an option for any future external payload mission where SR-71 transonic acceleration capability is questionable or in which test time at high Mach conditions must be maximized.
This paper presents the results of a preliminary analysis and simulation study of SR-71 performance with additional drag from externally mounted payloads. Maximum payload drag profiles with which the SR-71 aircraft can complete a research mission to the target test condition of Mach 3 at an altitude of 70,000 ft (21,336 m) with and without the use of the J58 thrust enhancement techniques are compared. In addition, the benefits of using thrust enhancement are shown using the drag profile for an air-launched GHRV.

A fuel-saving climb technique is discussed. The thrust ment options, along with the simulation methodologies this study, are also described.

enhanceused for

AIRCRAFT DESCRIPTION The SR-71 aircraft, built by Lockheed, is a long-range,
two-place, twin-engine airplane capable of cruising at speeds up to Mach 3.2 and altitudes up to 85,000 ft (25,908 m). 3 The aircraft is characterized by its black paint scheme; long, slender body; large delta wing; and prominent, spiked engine nacelles located midway out on each wing (fig. 3). Inboard and outboard elevons provide longitudinal and lateral control while twin, inboard-canted, all-moving rudders mounted on top of the nacelles provide directional control. The fuselage flares outboard along each side of the aircraft from the wing to the nose to form the chine. Within the nose

Length: 107 ft 5 in. Width: 55 ft 7 in. Minimum distance
between tails: 22 ft 9 in.
Aerial refueling

Forward (pilot's)

door -7 cockpit

Drag chute doors ter
flaps

Chine Nose section

Fuel tanks

Inlet spike

J58 engine

FIGURE 3. GENERAL ARRANGEMENT OF THE SR-71 AIRCRAFT.

and chine are several payload storage bays accessible from underneath the aircraft. The SR-71 aircraft would be structurally modified to carry large external payloads.
NASA Dryden currently has possession of two SR-71A vehicles and one B model. 4 The SR-71A models were previously used for military reconnaissance, whereas the SR-71B model was used for training. Except for its elevated second cockpit and nacelle ventral fins, the B model is similar to the A model in appearance and performance.

PROPULSION SYSTEM DESCRIPTION

The propulsion system of the SR-71 aircraft has three primary

components. These components are axisymmetric mixed compression inlets, Pratt & Whitney J58 turbojet engines, and

airframe-mounted, nozzles. 5

convergent-divergent

blow-in door ejector

The inlet spike (fig. 4) translates longitudinally, depending on Mach number, and controls the throat area. The spike provides ef-

ficient and stable inlet shock structure throughout the Mach range.

At the design cruise speed, most of the net propulsive force derives

from flow compression pressure on the forward facing surfaces of

the spike. Besides the spike, other inlet controls include the forward and aft bypass doors, used to maintain terminal shock position and to remove excess air from the inlet; and cowl and spike bleeds, used

to control boundary layer growth. The SR-71 aircraft is powered by two 34,000 lbf (151,240 N)
thrust-class J58 afterburning turbojet engines (fig. 5). Each engine

contains a nine-stage compressor driven by a two-stage turbine.

The main burner uses an eight-can combustor. The afterburner is

fully modulating. The primary nozzle area is variable. Above Mach 2.2, some of the airflow is bled from the fourth stage of

the compressor and dumped into the augmentor inlet through

six bleed-bypass tubes, circumventing the core of the engine and

transitioning the propulsive cycle from a pure turbojet to a

turbo-ramjet. The engine is hydromechanically

controlled and

burns a special low volatility jet fuel mixture known as JPT.

The inlet bleed and aft bypass flow mix with engine exhaust flow

just forward of the airframe-mounted ejector nozzle. Blow-in doors on the ejector nozzle remain open at low speeds and entrain additional mass flow into the exhaust stream. At high speeds, the doors

close and the airframe nozzle ejector flaps reposition to form a con-

vergent-divergent geometry. The blow-in doors and ejector flaps

are positioned by aerodynamic forces.

The engine spikes and forward bypass doors are positioned by commands from the digital automatic flight and inlet control sys-

tem (DAFICS). The DAFICS provides precise control of the terminal shock position. The DAFICS has significantly improved

vehicle performance and range and has virtually eliminated inlet

unstart, compared to the older analog control system.

THRUST ENHANCEMENT DESCRIPTION Several options for increasing the thrust of the J58 engine have
been analytically evaluated. 2 These options included increasing TT5; increasing compressor rotor speed; modifying the compressor bleed and inlet guide vane schedules; and increasing the augmentor fuel flow, WFAB, combined with oxidizer injection.
ErLolne Control Modifications
NASA Dryden desired that thrust enhancements allow for a minimum 50-hr engine life (time between overhaul). In response, Pratt & Whitney estimated that a combination of 150 RPM rotor speed increase and 75 °F (297 °K) TT5 increase would reduce the engine life to this 50-hr limit (fig. 6) while increasing net thrust an average of 5 percent throughout the Mach range of the vehicle (fig. 7). The second-stage turbine would be the life-limited component.

Aft bypass Spike bleed doors -

exit louvers ---_-'- _

Forward bypass

_'_L_

_---Translating_

\ spike _

\

/Y

\

/

\/

/_

/ ____._,-_

//_ \ _.........//_

Forward

__

bypass

_

Cowl bleed dOOrs

7

(shock trap)

z_Spike bleed

950042

FIGURE 4. INLET CUTAWAY VIEW WITH FEATURES.

FIGURE 5. THE J58 TURBOJET ENGINE.

E-25901

blade life,

hr 160_ I

Selected _tnms _ provide minimum

80 _- 50-hr life 7

increase, 0
25

0

1 O0

200

300

400

Rotor speed increase, RPM 950043

FIGURE 6. SECOND STAGE TURBINE BLADE LIFE LOSS AS A FUNCTION OF INCREASING ROTOR SPEED AND TURBINE EXHAUST TEMPERATURE.

+ 4 percent augmentor

15 fuel

flow inc_

10 --__2

o

Net thrust increase, percent

5 --_rease

"_-150

-5

I

.8

1.2

v
RPM rotor speed
III
1.6 2.0 2.4 Mach number

increase
II
2.8 3.2
95OO44

FIGURE 7. CUMULATIVE

NET THRUST

FROM INDIVIDUAL THRUST ENHANCEMENT

FOR 500 KEAS.

INCREASES OPTIONS

Increasing TT5 and rotor speed would involve relatively simple mechanical adjustments to the main fuel control that could be done in the field.
The augmentor fuel control would also require a mechanical adjustment that would increase WFAB by approximately 4 percent throughout the Mach range. This extra fuel flow would take advantage of the oxidizer available from the N20 injection, providing additional thrust.
The previous study did not recommend changing the bleed or inlet guide vane schedules without further analysis because of potential engine stability problems.

Nitrous Oxide In iection System
The previous study also concluded that augmentor oxidizer injection was feasible and recommended N20 as the oxidizer. 2 The N20 was suitable because of its ease in handling, nonvolatility, and lack of toxicity. Other oxidizers, including hydrogen peroxide, were ruled out primarily because of operational safety concerns or the need for extensive cryogenic handling procedures. The N20 injection, combined with a simultaneous WFAB increase, was predicted to provide at least an additional 5 percent net thrust increase throughout the SR-71 Mach range (fig. 7).
The study recommended that the N20 delivery system use four close-coupled spray rings, one residing within each aft-facing notch of the four standard flameholders. 2 Figure 8 shows a cutaway view of the forward augmentor duct along with the installation locations of the oxidizer spray rings. Quarter-inch, high-temperature flexline tubing comprising the spray rings would contain a total of 358 orifices, each 0.025 in. (0.635 mm) in diameter. The orifices would spray the N20 onto the inside surfaces of the flameholders.
Based on ground test results, this type of arrangement should provide good rmxing with the turbine exhaust air and effective decomposition of the oxidizer without causing thermal damage to the flameholders. Combustion efficiency is expected to remain as high as it is in a standard augmentor.
The previous study considered only the N20 delivery system. NASA Dryden developed a preliminary concept for storing the oxidizer within the vehicle. This application used portions of an SR-71 payload bay hydrogen storage concept developed for a proposed transonic drag reduction experiment in support of hypersonic vehicle research. 6 Storage tank volumes, weights, dimensions, and arrangements from that study were used to estimate a feasible payload bay N20 capacity range for the present application.
Additional oxidizer storage volume could also be traded for existing JP7 fuselage fuel tank capacity, but not without the very large expense required for such a major airframe conversion. This option was not thought to be feasible, so only the payload bay volume was considered for N20 storage.
Using data from the hydrogen storage design, it was estimated that a total volume of approximately 43 ft 3 (1.22 m 3) using 13 in-line high-pressure cylinders would be available for N20 storage. Figure 9 shows the proposed oxidizer tank arrangement in the SR-71 aircraft. A gaseous nitrogen purge system, along with connection and pressure regulation hardware, would fit comfortably in the remaining storage space. This system would provide a total liquid N20 mass capacity of approximately 2700 Ibm (1225 kg) at a storage temperature of 0 °F (255.4 °K) using a system pressure of less than 800 psia (5,516,000 N/m2). The payload bays are all supplied with cooling air from the environmental control system. The cooling air would be used to limit the temperature rise of the oxidizer before it is used. The total empty weight of the N20 storage and delivery system was estimated to be less than 2000 lb (907 kg).
The previous study suggested a metered N20 rate of 8 lb/sec (3.6 kg/sec) for each engine. With an on-board capacity of 2700 lb (1225 kg), the selected flow rate gives a total possible oxidizer delivery time of 169 sec. Higher flow rates are achievable, but not without risk of thermally choking the J58 nozzle. Regardless, the

/m

m

m

r

Nozzle

_ _-J/I

N20 sprayrings _li:hi:hstdnd:rd

__ / 11 J-_- I LJPTap_y "_'_:N _

_ /-Augmntor

-

_

_x__ High temperature--

flexline

950O45

FIGURE 8. AUGMENTOR CUTAWAY SHOWING PROPOSED N20 SPRAY RING LOCATION.

Top view

Nitrogen purge tank

N20 tanks (typical) -_/
Side view
95O046
FIGURE 9. N20 TANK ARRANGEMENT IN NOSE AND CHINE PAYLOAD BAYS.

increased thrust from extra flow would be offset by the disadvan-
tage of decreased delivery time. In combination, the engine control modifications and N20 injec-
tion were predicted to provide at least the l0 percent net thrust increase goal throughout most of the Mach range. The increase would be slightly less transonically.

DIGITAL PERFORMANCE

SIMULATION DESCRIPTION

NASA Dryden has developed a simplified aircraft performance simulation program that models many of its primary flight test vehicles. 7 This FORTRAN program, known as the digital performance simulation (DPS), consists of a main routine containing the

aircraft equations of motion and several subroutines. The subroutines include one for each of the modeled aircraft and contain the

lift, drag, thrust, and fuel flow data unique to that aircraft. A front-end subroutine creates an interactive interface from which the user can select the desired aircraft and then select which type of maneuver is to be modeled. Maneuver options include constant

Mach and airspeed climbs, level accelerations, and constant

g pushovers or pull-ups. The user is asked for initial and ending conditions. After execution, the DPS displays the output of the air-

craft states and performance as a function of time. The DPS equations of motion use four assumptions that simplify
the program while maintaining its fidelity for most maneuvers and applications: point-mass modeling, nonturbulent atmosphere, zero side forces, and a nonrotating Earth. The primary advantages of using the DPS over a piloted real-time simulator are that it is much easier to modify the aerodynamic and propulsion data tables, and

the DPS easily allows back-to-back comparisons of vehicle performance using a maneuver flown exactly the same in each case de-

spite a varying vehicle configuration. Also, the DPS computes aircraft performance in compressed time, allowing much faster viewing of the results than if the maneuvers were flown in real-time

on a piloted simulator. An SR-71 DPS module was developed specifically for this study
to predict vehicle performance with and without the drag and weight corresponding to the external payload and the additional thrust and fuel flow associated with the enhanced J58 engines. The

aircraft aerodynamic and propulsion data used to build the baseline

SR-71 model were supplied by Lockheed. 8 The thrust increase and fuel flow data for the engine enhancement options were supplied by Pratt & Whitney. Isolated hypersonic vehicle drag data were provided by the NASA Ames Research Center and were based on a derivative of the original GHRV. Lockheed supplied estimates of the

payload interference drag. The thrust enhancement data and payload drag data were then added to the DPS code as separate arrays and available as increments to the baseline data for use in the per-

formance computations. Once developed, the baseline DPS SR-71 option was verified

against flight data from the pilot's flight manual. 9 Different climb schedules were analyzed at standard day conditions. The resulting predicted elapsed time, fuel usage, and range data were compared to the values in the manual. The predictions compared closely to the actual flight data. It was believed the DPS model could be confidently used for estimating vehicle performance with additional

payload drag and the thrust enhancement options.

RESULTS AND DISCUSSION This section presents a fuel-saving climb technique for use with
a heavy payload during the ascent to the target test condition of Mach 3 at an altitude of 70,000 ft (21,336 m). Specific time and fuel savings using thrust enhancement with a high-drag payload mounted to the SR-71 aircraft are shown. The maximum external payload drag capability of the SR-71 aircraft using thrust enhancement is also presented.
O Dtimal Climb Schedule Selection
An early goal of this study was to determine the most fuel-efficient climb schedule for the SR-71 aircraft so that test time at the target test condition or the payload drag that could be accommodated to that condition would be maximized. This optimal schedule analysis focused on the transonic and higher speed range because the vehicle would be refueled at Mach 0.75 and an altitude of 25,000 ft (7620 m) soon after takeoff.
After refueling, the SR-71 aircraft is normally piloted along a constant Mach 0.9 climb to approximately 33,000 ft (10,058 m), pushed over at a 3000 ft/min (914 rn/min) descent to approximately 30,000 ft (9144 m), and pulled out in a level acceleration to 450 KEAS (Mach 1.25). This transonic penetration procedure is known as the climb-dive technique. At 450 KEAS, the vehicle initiates a constant equivalent airspeed climb to Mach 2.6. At this Mach number, equivalent airspeed is slowly reduced until the vehicle reaches the design cruise condition (3.2 maximum Mach number and 85,000 ft (25,908 m) maximum altitude). The climb-dive technique was discovered by Redin in the early 1970s to use less fuel than level transonic acceleration at 25,000 ft (7620 m) uses. 1°
Using high payload drag increments, the climb-dive technique was reevaluated using different pushover and pull-out altitudes to see if another combination gave greater fuel savings than were realized without payload drag in Redin's research. High pushover altitudes up to 37,000 ft (1t,278 m) offered large fuel savings. However, altitudes greater than 33,000 ft (10,058 m) were not used because, subsonically, the SR-71 aircraft can approach uncontrollable pitch-up conditions there. Pull-out altitudes below 28,000 ft (8534 m) offered slightly greater benefits than those above 28,000 ft (8534 m), but were avoided because focused sonic boom effects on the population become an issue.
The most feasible combination was a pushover altitude of 33,000 ft (10,058 m) combined with a pull-out altitude of 28,000 ft (8534 m). Also used was an increased descent rate of 6000 ft/min (1829 m/min) previously found by Redin to still offer a manageable piloting task. The selected schedule used approximately 2800 Ibm (1270 kg) less of JP7 than level transonic acceleration at 25,000 ft (7620 m) would have used. This fuel savings for the high-drag configuration was substantially higher than that previously predicted by Redin for a clean (no additional external drag) vehicle.
In the GHRV feasibility report, Lockheed recommended that the traditional 450 KEAS supersonic climb schedule be replaced with a 500 KEAS profile. The benefit of the higher KEAS schedule is an increase in excess thrust and the corresponding ability to handle higher payload drag than the 450 KEAS schedule can handle. This equivalent airspeed currently represents the right-hand edge of the conventional envelope and would require a very precise piloting technique to avoid overspeeding the aircraft. Otherwise,

Lockhedeiddnotforeseaenyoperabilpitryoblemflysingatthis equivaleanirtspeed.
Figure10showtshat,usingtheaboveselectetdransonic climb-divtechniquthee,timerequiretdoreachthetargetet st conditioanfterefuelinisgreducebdy4.1min(22percenuts)ing the500KEAS schedule compared to the 450 KEAS profile (no ex-
ternal payload drag in either case). A 3200 Ibm (1451 kg) savings in JP7 was also seen. These data, as with all data in the results sec-
tion, are presented for standard day atmospheric conditions. The 33,000/28,000 ft (10,058/8534 m) transonic climb-dive
technique and the 500 KEAS supersonic climb schedule were thus selected as offering the most fuel savings within the feasible operational limitations of the SR-71 aircraft. Figure 11 shows the resulting combined profile superimposed on the flight envelope and the maneuver sequences flown in the DPS to obtain this profile.

20--

/

16--

i/

Elapsed time, min

12 --

450 KEAS climb / /

Transonic 8

//j,,'_

-[.,_v=,_A=S bl;-;d-off

4 _ _500
o%1'1

.6

1.0

KEAS climb
II I I

1.4 1.8 2.2 Mach number

2.6 3.0
950047

Thrust Enhancement Benefits

With the climb schedule thus defined, the benefits of the thrust

enhancement options were quantified using an example external payload drag profile. The example profile was patterned after the GHRV. The modeled research vehicle was 47.2 ft (14.4 m) long and 17.1 ft (5.2 m) wide. The vehicle had a belly-mounted scramjet engine, one RL-10 liquid hydrogen rocket engine for boost, a cross-sectional area of 25.8 ft 2 (2.4 m2), and a total fueled weight of 13,700 Ibm (6214 kg). The vehicle was intended to be pylon-mounted on the back of the SR-71 aircraft between the vertical tails.

The maximum allowable weight of the SR-71 aircraft is approximately 140,000 Ibm (63,503 kg). The 13,700 Ibm (6214 kg) of external payload, 500 Ibm (227 kg) (estimated) of pylon structure weight, 2000 Ibm (907 kg) of oxidizer support hardware, and 60,000 Ibm (27,216 kg) of nominal dry vehicle weight leaves a total weight margin of approximately 64,000 Ibm (29,030 kg) available for JP7 and N20. After the external payload is air-launched at the target test condition, approximately 18,000 Ibm (8165 kg) of JP7 is required to safely land the aircraft. This value includes reserves if a postlaunch refueling attempt fails. The fuel control adjustments would not be selectable in flight and would provide thrust enhance-

ment during the entire mission. be controlled from the cockpit. of using the thrust enhancement of the DPS.

The N20 injection initiation would Using these constraints, the benefits
options were quantified by means

Figure 12 shows the two primary parameters used for judging mission feasibility: the N20 expended, and the JP7 remaining at the target test condition, both as a function of the N20 cutoff Mach

number. An unlimited N20 load capacity within the maximum weight limit of the vehicle is assurned. The in-flight refueling con-
dition of Mach 0.75 at an altitude of 25,000 ft (7620 m) is used as

the starting point. The N20 injection is initiated at Mach 1. This speed marks the beginning of the sharp drop in transonic excess thrust. Through analysis, it was shown that initiating oxidizer injection sooner in the acceleration offers little benefit in terms of over-

all fuel and time saved to reach the target test condition for launch. Figure 12(a) shows that the N20 can be injected until approxi-
mately Mach 1.23 before the supply is exhausted, assuming the maximum 2700 Ibm (1225 kg) load. However, figure 12(b) shows

FIGURE 10. ELAPSED TIME TO ACCELERATE TO MACH 3 AND AN ALTITUDE OF 70,000 FT (21,336 M) USING TWO DIFFERENT CLIMB SCHEDULES; NO ENHANCEMENT AND NO PAYLOAD USED.
that the JP7 remaining at the target test condition would be maximized with an N20 cutoff Mach number of 1.37. This cutoff Mach number is not achievable; however, there is only a slight penalty when comparing the JP7 remaining using a Mach 1.23 cutoff. Oxidizer injection produces thrust less efficiently than the standard engine, so the fuel-saving benefits of the N20 injection reverse once the vehicle pushes through the transonic drag rise and the excess thrust of the J58 engines rapidly increases.
The fuel remaining at the target test condition is approximately 21,400 Ibm (9707 kg) using the Mach 1.23 N20 cutoff. This cutoff Mach number was found to provide a total possible test time of nearly 4 min before the 18,000 Ibm (8165 kg) JP7 return-to-base limit is reached and the payload must be launched so that the SR-71 aircraft can begin its descent. Figure 12(b) also shows that using fuel control enhancements only (no injection--that is, an N20 cutoff of Mach 1), the SR-71 aircraft could not achieve the target test condition before reaching its return-to-base fuel value.
Figure 13 shows the time taken to reach the target test condition, after the completion of the in-flight refueling, as a function of the N20 cutoff Mach number. Figure 13 shows that much of the time-savings benefit comes early in the acceleration using oxidizer injection. In fact, the most benefit is obtained by expending the N20 within the transonic drag rise as opposed to delaying N20 delivery until later in the acceleration.
Figure 14 shows comparisons of elapsed time as a function of Mach number for the SR-71 climb with and without the thrust enhancement and payload drag increments. For each case, the vehicle begins the acceleration at the maximum allowable vehicle weight of 140,000 Ibm (63,503 kg). Curve 1 shows the baseline characteristics of the clean vehicle with no thrust enhancement. Curve 2 represents the vehicle with the additional payload drag and with both fuel control and oxidizer thrust enhancements active. The N20 was initiated at Mach 1 and used until fully expended. Despite the extra propulsive force from thrust enhancement, the vehicle still takes an

90 x 10 3

80 -70 60

450 KEAS (standard
climb path)--_..""

..i

..... °.. - -"°"".... i,

.-"'" .-'"

O
_'° . f7

j_/

AItitude,ft

__

.."°'"'" __

40 -- .."_""

," _ 30 _current

,i //

/
L 500 KEAS operating limit)

Climb schedule with large payload
1) Level acceleration at 25K ft at
minimum augmentation
2) Constant Mach 0.9 climb at
maximum augmentation
3) Pushover to 6000 ft/min descent
from 33K ft
4) Pull-out at 28K ft and subsequent
level acceleration
5) Initiate 500 KEAS climb at Mach 1.32 S) Begin linear decrease in KEAS at
Mach 2.6
7) On condition at Mach 3 and 70K ft
(416 KEAS)

20 _'1°-'021 ./

i

I

[

I

.6

1.0

1.4

1.8

2.2

2.6

Mach number

I;

3.0

3.4

_oo4,

FIGURE 11. THE SR-71 FLIGHT ENVELOPE WITH RECOMMENDED 500 KEAS CLIMB SCHEDULE PROCEDURE.

16 x 103

14-12--

N20 expended,
Ib

10 -86-4

0 1.0

S

bay storage limit

IIIII

1.4

1.8

2.2

2.6

3.0

N20 cutoff Mach number 9soo49

(A) N20 EXPENDED AT MACH 3 AND AN ALTITUDE OF 70,000 FT (21,336 M).

JP7 remaining,
Ib

23x 103
22 21 .

Mach 1.23 maximum speed with 2.7K Ib
. ,2o,o.d

cutoff

I
18
17
16 1.0

18K Ib retu rn-to-base

JP7 limit
II

1.4

1.8

2.2

%
" %

II

2.6

3.0

N20 cutoff Mach number 950050

(B) JP7 REMAINING AT MACH 3 AND AN ALTITUDE OF 70,000 FT (21,336 M).

FIGURE 12. N20 EXPENDED FUNCTION OF N20 CUTOFF PAYLOAD USED.

AND JP7 REMAINING AT MACH 3 AND MACH NUMBER; N20 INJECTION INITIATED

AN ALTITUDE OF 70,000 FT (21,336 M) AS A AT MACH 1; 500 KEAS CLIMB SCHEDULE AND

10

28 26 24 Time,22
min 20

Mach 1.23 maximum cutoff speed with 2.7K Ib N20 load

18--
III1

1.4

1.8

2.2

2.6

N20 cutoff Math number

3.0 9500_1

FIGURE 13. TIME TO REACH MACH 3 AND AN ALTITUDE OF 25,000 FT (7620 M) AS A FUNCTION OF N20 CUTOFF USED.

OF 70,000 FT (21,336 M) FROM MACH 0.75 AND AN ALTITUDE MACH NUMBER; 500 KEAS CLIMB SCHEDULE AND PAYLOAD

28 24

20

Elapsed time, min

m
16
12

s,_ _ s

•_

s_ ,'_

..3
o s•

Payload?

1

N

2

Y

3

Y

4

N

5

N

Fuel control
enhance? N Y Y Y Y

2.7K Ib N20
enhance? N Y N N Y

8

4

.JE I I I I I

O6

1.0

1.4

1.8

2.2

2.6

3.0

Mach number

FIGURE 14. ELAPSED TIME TO ACCELERATE TO MACH 3 AND AN ALTITUDE OF 70,000 FT (21,336 M) USING VARIOUS PAYLOAD AND ENHANCEMENT OPTIONS; 500 KEAS CLIMB SCHEDULE USED.

extra 8.2 min (56 percent longer) to reach the target test condition because of the payload. Using fuel control enhancements only, in

addition to the payload, the vehicle barely accelerates through the drag rise, taking 84 percent longer (curve 3) than in the baseline

case. The SR-71/payload combination could not penetrate the drag

rise at all with all thrust enhancements curve is not shown.

disabled, so the resulting

Figure 14 also shows the results using thrust enhancement with no extra payload drag. Curve 4 shows that, using fuel control en-

hancement only, acceleration time is reduced 14 percent, whereas if the N20 injection is also used, acceleration time is reduced by

19 percent (curve 5). These last two curves show that for little or no additional payload drag, only small benefit can be gained from the thrust enhancement techniques. However, curves 2 and 3 indicate that for large additional payload drag, thrust enhancement is essential.
Maximum External Payload Drag CaDability
Figure 15 shows representative maximum vehicle/payload drag profiles that the SR-71 aircraft can accommodate with and without using thrust enhancement. Using these drag profiles, the vehicle

11

45x10 4O

3 Clean vehicle drag

/ • .....

Max, no enhancement

. m Max, full enhancement

35
Drag, Ibf 30
25

Basic vehicle capability

2O
is
1.0

1

I

t

1.4

1.8

2.2

Mach number

'I

2.6

3.0

950053

FIGURE 15. MAXIMUM PROFILES, WITH AND MENT, 4 MIN CRUISE SCHEDULE USED.

TOLERABLE PAYLOAD DRAG WITHOUT THRUST ENHANCEAT MACH 3; 500 KEAS CLIMB

would have nearly 4 min of fuel remaining at the target test condition of Mach 3 at an altitude of 70,000 ft (21,336 m) before having to initiate a return to base. This time value represents an arbitrary test-point duration.
Figure 15 shows that the maximum tolerable drag profile using thrust enhancement equates to an increase in drag capability of approximately 25 percent over the capability using no enhancement, both relative to the clean SR-71 drag profile (also shown). Once through the transonic drag rise, the absolute drag decreases as the vehicle gains altitude.
It was previously shown that the SR-71 aircraft, using thrust enhancement, would be able to complete the same approximate 4-min mission window with the additional drag from the hypersonic research vehicle.
Recommendations
If an external payload package idea is carried beyond the design phase and becomes a serious candidate for mounting on the SR-71 aircraft, empirical data on the drag characteristics of the SR-71/payload combination should be obtained before development commences, particularly if the payload is large or is a high-drag design. The empirical data would most likely be obtained in wind-tunnel testing because a large subscale test model of the SR-71 aircraft is available. The transonic drag rise traits would be of particular interest in such an analysis.
The empirical drag data would then be implemented in the high-fidelity SR-71 piloted simulator at NASA Dryden to verify that the vehicle could accommodate the payload on the desired mission from a performance standpoint. Other payload effects, particularly stability and control characteristics, would need to be

evaluated simultaneously. If the vehicle could not adequately complete the desired mission, it would then be necessary to implement the thrust enhancement techniques.
These techniques also would have to be evaluated empirically. A probable scenario is that one J58 engine would be modified and evaluated in ground testing to verify predicted performance increases. Then the modifications would be evaluated in flight, using a short-duration N20 injection sequence. The actual engine performance data would then be included in the piloted simulation and the mission reevaluated for feasibility. If engine operability and
performance are as predicted and aircraft performance is acceptable, the second engine would be modified and the full set of N20 hardware installed.

CONCLUDING REMARKS

The NASA Dryden Flight Research Center evaluated SR-71 per-

formance for the ability to accommodate large external payload

drag profiles with and without the use of engine

thrust-enhancement

techniques. The evaluation was performed

using a simplified aircraft performance simulation.

A previous study concluded that fuel control modifications and

nitrous oxide injection into the augmentor can feasibly increase the

net thrust of the J58 turbojet engine by an average of 10 percent

throughout the transonic and supersonic Mach range. The most cost-effective oxidizer storage scenario would use the nose and chine payload bays of the vehicle for a maximum capacity of 2700 Ibm (1225 kg) of nitrous oxide stored at 0 °F (255.4 °K). The extra thrust produced by the control modifications and transonic oxidizer injection would allow for an increase in external payload drag capability of approximately 25 percent.
The SR-71 flight profile was evaluated to determine an optimal climb sequence for use with a high-drag payload. Of particular interest was the transonic speed range and associated drag rise. The most feasible profile used a maximum power, Mach 0.9 climb from an altitude of 25,000 ft (7620 m) to an altitude of 33,000 ft (10,058 m); a pushover and subsequent 6000 ft/min (1829 m/min) descent to an altitude of 28,000 ft (8534 m); a level acceleration to 500 KEAS; and then a constant 500 KEAS climb to the high Mach number target test condition of Mach 3 and an altitude of 70,000 ft (21,336 m). This maneuver was found to save nearly 6000 Ibm (2721 kg) of jet fuel over the more conservative transonic level acceleration and 450 KEAS climb schedule.

The performance of the SR-71 aircraft was evaluated using a payload drag profile example representative of a large hypersonic research vehicle designed for high-speed air-launch. Without the use of thrust enhancement, the vehicle could not accelerate through

the transonic range. Using fuel control modifications only, the vehicle was able to accelerate through the transonic range but was unable to reach the target test condition before reaching the return-to-base jet fuel limit. Using nitrous oxide injection from Mach 1 to Mach 1.23, which fully depleted the 2700 Ibm (1225 kg) load, the vehicle was able to reach the target test condition with ap-

proximately 4 min of cruise time available before having to return to base.
The results show that the SR-71 aircraft has a large payload drag

capability even without the use of thrust enhancement. However,

12

forvoluminopuasyloaodrspayloatdhsaat pproatchhe20,00I0bm (907k2g)externlaoladcapaciotyf thevehicleit,ispossibtlheat thepayloaddragwillexceethdeperformacnacpeabiluitynlesthse thrusetnhancemteecnhtniquaeresused.

REFERENCES

1Hicks, John W., Flight Testing of Airbreathing Vehicles, NASA TM-4524, 1993.

Hypersonic

2United Technologies Pratt & Whitney, Government Engines

and Space Propulsion Division, J58 Thrust Upgrade Study for

Hypersonic Air Launch Option Concept, FR 22741-1, Mar. 1993.

3Miller, Jay, Lockheed SR-71 (A.12/YF-12/D-21), Arlington, Texas, 1985.
4Lockheed Advanced Development Company,

Aerofax, Inc., Lockheed SR-7I

Supersonic/ltypersonic

Research Facility. Researcher's

book; Volume 11: Technical Description, 1992.

Hand-

5Burcham, Frank W., Jr., Montoya, Earl J., and Lutschg,

Phillip J., Description Instrumentation for TM-X-3099, 1974.

of YF-12C Airplane, Propulsion System, and Propulsion Research Flight Tests, NASA

6Hicks, John W., "Development of a Hydrogen External Burning Flight Test Experiment on the NASA Dryden SR-71A Aircraft," SAE-92-0997, Apr. 1992.
7Orme, John S., Digital Performance Simulation Models of the F-15, F-16XL F-18, F-104, TACT F-111, X-29 and Hypersonic Research Vehicle, NASA TM- 104244, 1992.

8Meyer, J.E., McMaster, J.R., and Moody, R.L., Handling Qual-

ities of the SR-71 (Revised), Lockheed Aircraft Corporation SP-508, Sep. 1978.
9Lockheed Aeronautical Systems Company, SR-71A Flight Manual, SR-71A-I, Issue E, Change 2, Jul. 1989.

I°Redin, Paul C., Optimization of Transonic Acceleration Performance for the YF-12C Airplane Using the Climb-Dive

Maneuver, TM-X-2694, 1973.

13

REPORT DOCUMENTATION

PAGE

Form Approved OMB NO. 0704-0188

Pubt_c reporting burpen for this collection of information is estimated to average 1 hour per response, including the time for reviewing instruCtions,

searching existing data sources, gathering and

maintaintng the data needed, and completing

and reviewing the collection of information. Send comments regarding this burden estimate or any other aspect of this collection of informat=on,

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for reducing this burden, to Washington Hee0quarters

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and to the Office of Management

and Budget, Paperwork

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Reduction Pro_ect (0704o0188).

Washington,

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1215 Jefferson

Davis Highway,

Suite 1204, Arhngton.

1. AGENCY USE ONLY (Leave blank)

2. REPORT DATE
June 1997

3. REPORTTYPE AND DATES COVERED
Technical Memorandum

4.TITLE AND SUBTITLE
Predicted Performance External Payload
6. AUTHOR(S)

of a Thrust-Enhanced

SR-71 Aircraft with an

5. FUNDING NUMBERS
WU 242 33 02 00 23 00T15

Timothy R. Conners

7.PERFORMINOGRGANIZATIONANME(SA)NDADDRESS(ES)
NASA Dryden Flight Research Center P.O. Box 273 Edwards, California 93523-0273

8. PERFORMING ORGANIZATION REPORT NUMBER
H-2179

9.SPONSORING/MONITOARGINEGNCNYAME(SA)NDADDRESS(ES)
National Aeronautics and Space Administration Washington, DC 20546-0001

10. SPONSORING/MONITORING AGENCY REPORT NUMBER
NASA TM- 104330

11. SUPPLEMENTARY NOTES
Presented at the International June 5-8, 1995.

Gas Turbine and Aeroengine

Congress

and Exposition,

Houston, Texas,

12a. DISTRIBUTION/AVAILABILITY

STATEMENT

12b. DISTRIBUTION CODE

Unclassified--Unlimited Subject Category 05

13. ABSTRACT (Mulmum 200 words)
NASA Dryden Flight Research Center has completed a preliminary performance analysis of the SR-71 aircraft for use as a launch platform for high-speed research vehicles and for carrying captive experimental packages to high altitude and Mach number conditions. Externally mounted research platforms can significantly increase drag, limiting test time and, in extreme cases, prohibiting penetration through the high-drag, transonic flight regime.
To provide supplemental SR-71 acceleration, methods have been developed that could increase the thrust of the J58 turbojet engines. These methods include temperature and speed increases and augmentor nitrous oxide injection. The thrust-enhanced engines would allow the SR-71 aircraft to carry higher drag research platforms than it could without enhancement.
This paper presents predicted SR-71 performance with and without enhanced engines. A modified climb-dive technique is shown to reduce fuel consumption when flying through the transonic flight regime with a large external payload. Estimates are included of the maximum platform drag profiles with which the aircraft could still complete a high-speed research mission. In this case, enhancement was found to increase the SR-71 payload drag capability by 25 percent. The thrust enhancement techniques and performance prediction methodology are described.

14. SUBJECTTERMS
Aircraft performance, J58 turbojet engine, Optimal flight profile, SR-71 aircraft, Thrust enhancement, Transonic drag rise

17. SECURITY CLASSIFICATION OF REPORT
Unclassified

18. SECURITY CLASSIFICATION OFTHIS PAGE
Unclassified

19. SECURITY CLASSIRCATION OF ABSTRACT
Unclassified

NSN 7540-01-280-5500

Available from the NASA Center for AeroSpace Information, 800 Elkridge Landing Road, L inthicum Heights. MD 21090, (301)621-0390

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17
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At3
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Unlimited
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