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Full text of "NASA Technical Reports Server (NTRS) 19990009345: The Role of the Strutjet Engine in New Global and Space Markets"

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THE ROLE OF THE STRUT JET ENGINE IN NEW 
GLOBAL AND SPACE MARKETS 


A . Siebenhaar and M. J. Bulman 
GenCorp Aerojet 
PO Box 1322 
Sacramento, California 


D. K. Bonnar 
Boeing Phantom Works 
Huntington Beach, California 


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List Of Tables 


In Text 
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Title 

Strut Rocket Design Parameters 
Evaluation Ranges of Subscale Strutrocket Injector 
Degree of Technology Relative To TRL 6 
Baseline Missions Selected for Rapid Cargo Delivery 


Reference 

Table 7 AIAA 
Table 8 AIAA 
Table 1 1 AIAA 
Table 1, Dave Bonnar 


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List Of Figures 


In Text 
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Title 

Strutjet Concept Provides Engine Robustness - The 
Prerequisite for Lower Cost for Access to Space and 
Global Reach Markets 

Strutjet Engine Has Clean Unobstructed Flowpath and 

Simple 2-D for Inlet and Nozzle Variable Geometry 

Strutjet Engine Concept 

Simple 2-D Variable Geometry Inlet Allows for 

Geometrical Contraction Variation 

Strutjet Engine Cycles and Operation 

Wave Rider-Type Mach 8 Cruise Missile Provides 

Significant Forebody Compression 

Axisymmetric Mach 8 Strutjet Engine Facilitates 

Engine-Vehicle Integration and Provides Centerline 

Thrust Vector 

Subscale Strutjet Inlet Test Hardware Mounted In 
Wind Tunnel 

Subscale Strutjet Inlet Test Hardware 

First Strutjet Inlet Generated High Pressure Ratios 

Missile Size Strutrockets 

Strutrocket Chamber Geometry and Propellant Feeds 
For Test Article 

Strutjet Uncooled Test Engine With Adjustable 2-D 
Combustor Geometry 

Direct Connect Ducted Rocket Test Data Shows Peak 
Thrust Enhancement of 13% 

Fuel Rich Rocket Doubles Thrust In Ducted Rocket 
Mode 

Both Gaseous Ethene and Pilot Vaporized Cold JP-10 
Yield High Combustion Efficiency 
Piloted Fuel Injection Essential For Ignition and 
Sustaining Of Combustion At Mach 4 
Mach 7 Combustion Achieved High Efficiency With 
Rapidly Expanding Geometry 
Freejet Engine Installed In Wind Tunnel 
Good Agreement Obtained Between Inlet Only and 
Freejet Internal Pressure Measurements 
Net Thrust Increase Versus Fuel Flow in Freejet Tests 
Strutjet Inlet In Wind Tunnel 
Test Data Exhibits Large Unstart Margin 
Single Rocket Chamber Ignition Test 
Individual Direct Connect Strut 
Two Struts In Direct Connect Rig 
Cascade Injectors Integrated In Test Hardware 
Strutrocket Integrated Into Strut Base 
Frontal View Of Freejet Engine (Scale Intentionally 
Distorted) 

Aft View Of Freejet Engine (Scale Intentionally 
Distorted) 

Internal View of Freejet Engine (Scale Intentionally 
Distorted) 


Reference 
Fig. 24 AIAA 


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32 

Primary Cargo Missions In The Pacific Rim 

From Dave Bonnar 

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RBCC Engine Model Shows High Specific Impulse 
During Airbreathing Modes 

From Dave Bonnar 

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Vehicle/Engine Drag Reduces Specific Impulse 

From Dave Bonnar 

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RBCC Engine Thrust Model 

From Dave Bonnar 

36 

RBCC Propane Vehicle For Cargo Missions Similar 
Size As Space Shuttle Orbiter 

From Dave Bonnar 

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Ground Launched Vehicle Design - RBCC/HTHL 

From Dave Bonnar 

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RCDS-GL Vehicle Ignition Mass 

From Dave Bonnar 

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RCDS-GL Flyout Trajectory 

From Dave Bonnar 

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RCDS-GL Maximum Range 

From Dave Bonnar 

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RCDS-GL Dynamic Pressure Profile 

From Dave Bonnar 

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RBCC Mission Range Decreases With Lower 
Maximum Heat Flux 

From Dave Bonnar 

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Heat Flux Profile 

From Dave Bonnar 

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Temperature Profile 

From Dave Bonnar 

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RCDS-GL Vehicle Mission Propellant Weight 
Requirements 

From Dave Bonnar 

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Air Launched Vehicle Design - RBCC/HTHL 

From Dave Bonnar 

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AN-225 Air Launch Platform for RBCC/ HTHL 
Vehicle 

From Dave Bonnar 

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RCDS-AL Vehicle Falls Within Launch Aircraft 
Capability 

From Dave Bonnar 

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RCDS-AL Vehicle Shows Burnout Altitude At 180 kft 
for Maximum Range 

From Dave Bonnar 

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RCDS-AL Vehicle Flies From Anchorage To Beyond 
Brisbane To Perth 

From Dave Bonnar 

51 

RCDS-AL Vehicle Mission Propellant Weight 
Requirements 



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Abstract 

The Strutjet, discussed in previous IAF papers, was originally introduced as an enabling 
propulsion concept for single stage to orbit applications. Recent design considerations 
indicate that this systems also provides benefits supportive of other commercial non- 
space applications. This paper describes the technical progress of the Strutjet since 1997 
together with a rationale why Rocket Based Combined Cycle Engines in general, and the 
Strutjet in particular, lend themselves uniquely to systems having the ability to expand 
current space and open new global "rapid delivery" markets. 

During this decade, Strutjet technology has been evaluated in over 1000 tests. Its design 
maturity has been continuously improved and desired features, like simple variable 
geometry and low drag flowpath resulting in high performance, have been verified. In 
addition, data is now available which allows the designer, who is challenged to maximize 
system operability and economic feasibility, to choose between hydrogen or hydrocarbon 
fuels for a variety of application. The ability exists now to apply this propulsion system to 
various vehicles with a multitude of missions. 

In this paper, storable hydrocarbon and gaseous hydrogen Strutjet RBCC test data as 
accomplished to date and as planned for the future is presented, and the degree of 
required technology maturity achieved so far is assessed. Two vehicles, using cryogenic 
propane fuel Strutjet engines, and specifically designed for rapid point-to-point cargo 
delivery between Pacific rim locations are introduced, discussed, and compared. 


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1.0 Introduction 

The premise of this paper is: “Low Cost Transportation Will Enable New Markets”, or 
new global and space markets can be opened if government and industry leaders 
subscribe to a change in the decades-old mind set about what is “necessary” to operate 
launch vehicles 1 . 

Let’s look at system development cost. The shuttle orbiter took about $12 billion to 
develop, and at eight, or fewer, flights per year the cost to launch payloads amounts to 
$10,000/lbm just to pay for the development cost. In contrast, a new commercial aircraft, 
like the Boeing 777, with a development cost of $7 billion amortizes this investment over 
its lifetime with a payload cost of about $10/lbm. Obviously, the former is not a viable 
basis for business, and the latter is very good business. When looking at the systems 
operating cost the situation does not improve. 

One would assume that today’s reusable launch vehicle community (RLV) would endorse 
the airline approach, but the fact is that most commercial reusable vehicles are still being 
built more like space shuttle orbiters than commercial aircraft. In order to bring the cost 
down, these vehicle are designed for a life of 50 to 100 flights, too meager an 
improvement, because at the optimistically projected payload cost of $1000/lbm the 
production cost can just be covered, and a higher price must be charged to cover 
operating cost and profits. If these systems could be designed for, let’s say, 1000 to 2000 
flights per life, the payload cost could be reduced to $10 to 100/lbm, and new rapid 
global reach and orbit markets could be realized 2 . 

The key to such a capability is the abandonment of the extremely high power density of 
the classical rocket as the sole means of propulsion. Low cost systems need engines and 
subsystems capable of operating for thousands of hours between overhaul, have built-in 
self-checking capabilities, and are designed for quick replacement in the field without 
disturbing other systems. The Aerojet Strutjet engine 3, 4 is an engine with the potential to 
provide simultaneously higher performance, more safety, and higher reliability than 
rockets. An investment in its high-reliability design will quickly pay off in faster turn- 
around times, higher utilization of each vehicle, customer confidence, and greatly reduced 
insurance rates. Traditionally, rockets must be built with small margins because 


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inherently high mass fraction rocket vehicles require all the performance achievable. This 
strategy is acceptable for expendable systems, but it leads to a major downfall in the 
attempt to adopt it to reusable systems. Here robustness, obtained in exchange for 
performance, is the key parameter from which all desirable attributes of low cost through 
reusability can be derived. The point in case is illustrated for space access missions in 
Figure 1 which relates lower engine replacement cost to the 80% reduced thrust 
requirement, and lower maintenance and operating cost to higher structural margins and 
redundancy. Both these savings result from the increased specific impulse of RBCC 
engines and the associated reduction in takeoff and dry mass. 

Future markets lie beyond space access. If RLV technology is brought to bear to 
hypersonic transport of rapid delivery systems for first cargo and possibly later 
passengers between key traffic centers, the RLV traffic will jump to thousands of flights 
per year. This will require a more airport-like layout of launch operations including noise 
mitigation, air-traffic control, and, because of high acquisition and storage cost, a 
departure from liquid hydrogen as the main fuel. Hydrocarbon fuels, e.g. cryogenic 
methane or propane, or other kerosene based jet fuels will be the propellants of choice. 
The key to successful business is a vehicle/operations/infrastructure system design 
focusing from the outset on the efficient achievement of high flight rates. Structuring the 
system to expand rapidly to fill new markets will have a large economic advantage over 
today’s approach with systems designed for traditional launch rates and subsequently 
applied to a demand-elastic market. 

This paper describes briefly in Section 2 the space access version of the Strutjet engine, 
and provides in Section 3 a status of Strutjet RBCC test data addressing both storable 
hydrocarbon and gaseous hydrogen fuels. In Section 4 an assessment of the current 
maturity level of required Strutjet technology levels is made. Finally, Section 5 
introduces a preliminary study of an rapid cargo delivery system capable of delivering 
payloads from the United States to various locations on the Pacific rim. This study is to 
be viewed as an appetizer to stimulate future further exploration of these potential 
capabilities of RBCC propulsion. Section 6 sums up the overall paper content, and 
Section 7 provides the relevant references. 


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2.0 Space Access Strutjet Engine Description 

The Aerojet Strutjet engine is a member of the rocket based combined cycle 
(RBCC) class of engines with several new technologies and innovations. Many of these 
technologies are also of far reaching interest in other propulsion schemes. The Strutjet 
engine operates all of the cycles with liquid hydrogen (LH2) for fuel, and liquid oxygen 
(LOX) and atmospheric air as necessary for combustion of fuel. 

The name Strutjet is derived from the use of a series of struts in the front part of 
the engine flowpath. The struts serve a number of functions in the engine, including 
compression of incoming air, isolation of combustion from air inlet (that is, as an 
isolator), fuel distribution and injection, ram/scram combustion, and rocket-thruster 
integration in the different modes of engine operation. The struts provide efficient 
structural support for the engine. The struts thus form a key element in the engine. 

The Strutjet engine is in principle, a single engine configuration with three 
propulsion elements, namely rocket, ramjet, and scramjet in its five mode operation. The 
elements are highly integrated in design and function. The Strutjet differs from other 
RBCC designs in the higher degree of functional integration of various engine 
components which results in a shorter, higher thrust-to-weight engine with good 
performance. The specific impulse of the Strutjet engine is characteristic of other known 
airbreathing engines without the thrust-to-weight penalty of having separate propulsion 
systems for different flight conditions. 

The Strutjet engine concept is founded on obtaining specific impulse, I S p , higher 
than that of a rocket, a better thrust-to-weight ratio (F/W e ) than an all-airbreathing 

engine, and a substantial reduction in vehicle gross take-off-weight. Thus along a typical 
trajectory of a single stage to orbit (SSTO) vehicle the mean I S p for the Strutjet is 

estimated to be 586 sec, while the conventional hydrogen-oxygen rocket provides I S p of 
only 425 sec. In comparison, the mean I S p for an all-airbreather such as the USA 


S 



National AeroSpace Plane (NASP) was estimated to be 755 sec. A SSTO-type all-rocket 
may yield FAV g equal to 80:1 and a NASP-type SSTO all-airbreather, about 6:1. Using 

current state-of-the-art technology, a hydrogen fueled Strutjet can be built with F/W e of 

35:1; however, for the sake of increased engine robustness, low earth orbit missions can 
still be achieved with F/W values as low as 25: 1 . 

c 

Figure 2 provides a schematic of the Strutjet and identifies the different elements 
in the combined cycle engine. The propulsion subsystems are integrated into a single 
engine using common propellant feed lines, cooling systems, and controls. The air inlet 
along with the struts, the combustion sections, and the nozzle make up the main engine 
flow path. An isometric view of a typical RBCC Strutjet engine is given in Figure 3. 

The variable geometry inlet incorporates two engine ramps, that maximize air 
capture and control compression as required by the engine. The inlet combines effective 
forebody precompression with strut compression. This results in "soft start", low spill 
drag, and good capture and recovery efficiencies. The "soft start" is a result of the 
increased openness of the inlet on the cowl side, which causes a gradual decrease in 
spillage with increasing Mach number. This, in turn, provides smooth increases in 
captured air mass flow and pressure recovery. The inlet geometry and the changes in 
contraction ratio as a function of flight Mach number are shown in Figure 4. Past the 
cowl lip, at low speeds, the flow area between two adjacent struts remains constant; the 
strut section serves as an inlet combustor isolator during ducted rocket and ramjet mode. 
At higher speeds (M«5) the inlet ramps are deployed to increase the engine contraction 
and performance. 

Above Mach 6 the scram mode is the most efficient and the diverging isolator duct 
between the fully contracted inlet throat and the rockets is used as the scram combustor. 
During this mode of operation the variable nozzle geometry adjusts the scram combustor 
flowpath into a continuously diverging configuration. 


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The ram combustor used at speeds up to Mach 6, is located aft of the strut rockets. This 
combustor provides enough area to permit stoichiometric subsonic combustion at low 
speeds. 

The Strutjet variable geometry nozzle is a simple flap used to control subsonic 
combustion pressure to create the optimum thrust as dictated by operating mode and 
flight Mach number. This nozzle flap is the principle control in the transition to the 
scram mode. By opening the nozzle flap up at approximately Mach 6, the combustor 
pressure drops and the flow remains supersonic through the combustor. 

The overall Strutjet engine propellant flow is illustrated in Figure 5. As shown, 
there are three subsystems: (i) the hydrogen and oxygen fuel tanks, turbopumps feed 
system and powerhead; (ii) the strutrocket and fuel injection assembly, and (iii) the 
engine structure and cooling system. Both the strut rocket and the engine structure are 
operating in the thermal and combustion gasdynamic environment. 

Fuel rich gases generated in the fuel gas generator (FGG) drive the hydrogen 
turbine, and are subsequently injected into the engine internal air stream at selected 
locations through base-axial, aft, and forward injectors. The selection depends on the 
engine operating mode, and is accomplished through appropriate valving. Hydrogen gas 
is also used to cool the rocket chambers (the figure including schematically only one) and 
the engine structure before injection into the combustor section. In the expander cycle, 
hydrogen heated by the engine structure bypasses the prebumer and drives the turbine in 
an expander cycle. 

The oxygen side of the propellant system is only active during rocket operation, 
and operates (always) in a stage-combustion cycle. The oxygen rich turbine drive gases 
are generated in the prebumer (OPB) through the burning of a small amount of hydrogen 
and all of the oxygen flow, prior to injection into the rocket chambers. 


10 



The table shown at the bottom in Figure 5 includes four attributes for each of the five 
operating modes, and indicates graphically the required settings for a particular mode: 

(i) selected power cycle in the oxygen and hydrogen supplied, 

(ii) amount of rocket propellant flow, 

(iii) amount of hydrogen injected into the air stream, and 

(iv) settings of the inlet and nozzle variable geometry. 


Along a typical SSTO ascent trajectory a Strutjet operates in five modes with 
smooth transition between modes: 

I. ducted rocket operation for takeoff and acceleration through the transonic 
speed regime into the supersonic region; 

II. ramjet operation from Mach 2.5 to about 6; 

III. scramjet operation from Mach 6 into the hypersonic speed range up to Mach 

10 ; 

IV. scram /rocket operation from Mach 1 0 to low vacuum conditions; and 

V. ascent rocket from low vacuum operation up to orbital speeds. 

3.0 Available Hydrocarbon and Hydrogen Test Data and Planned Future Test 
Activities 

The overall test program, accomplished to-date or planned for the near future, may be 
divided into two groups: (i) missile propulsion tests using storable hydrocarbon 

fuels, and 

(ii) space launch propulsion tests using gaseous hydrogen fuel. 

3.1 Storable Hydrocarbon System Tests 


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Aerojet performed a test program which parametrically examined the RBCC Strutjet 
propulsion system from Mach 0 to 8 over the altitude range from 0 to 100,000 ft. The 
tests were carried out in the context of a long range missile, two configurations of which 
are shown in Figures 6 and 7. However, the test results obtained during this campaign are 
equally applicable to a launch system Strujet like the one described in Section 2. In over 
1000 hot fire & inlet tests a number of achievements were realized: 

• The strut inlet provides excellent air capture, pressure recovery, and unstart margin. 

• The integration of compact high chamber pressure rockets using gelled hypergolic & 
cryogenic propellants into a strut is structurally and thermally feasible. 

• A fixed engine flowpath geometry suitable for all modes of operation can be 
established providing adequate thrust and specific impulse to accomplish mission 
objectives. 

• Static sea level thrust augmentation of 13% can be achieved due to the interaction of 
air ingested with the fuel-rich rocket plume. 

• The ducted rocket thrust increases with increased flight Math number. At Mach 2.85 
and altitude 20,000 ft the thrust increase is over 100%. And, at Mach 3.9 and altitude 
40,000 ft. the ramjet thrust exceeds the rocket sea-level thrust by 19%. 

• Dual-mode operation of the ram-scram combustor with a thermally choked nozzle is 
feasible. 

• Efficient combustion at high altitude and with short combustors is possible with 
hypergolic pilots. Combustion efficiency of 90% can be demonstrated with a 
combustor only 30 in. long at Mach 8 conditions. 

Inlet Tests 

As part of the first Strutjet inlet test program, a subscale inlet model, shown in Figures 8 
and 9, was constructed to evaluate design options for the freejet engine inlet design. Prior 
to this testing, engine/vehicle performance was based on extrapolating the performance of 
the inlet from the literature (mostly the work of NASA LaRC). The objectives of the 
inlet development was to define a missile like inlet that would interface with the 
hydrocarbon combustor which up to this point had only been tested in a direct connect 


12 



configuration. A very conservative design was selected that had no internal contraction to 
assure starting. This inlet was found to start at all tested Mach numbers (4-6), and as 
shown in Figure 10, it produced excellent pressure rise. As predicted, it exhibited low 
energy flow near the body side of the flowpath, a phenomenon which had to be dealt with 
in the subsequent ffeejet tests via matching the fuel injection with the actual air flow 
distribution. 

Ducted Rocket Tests 

In these tests, each strut contained three water-cooled gelled IRFNA and MMH 
propellant rockets, as shown in Figure 11. The injector pattern consisted of 36 pairs of 
fuel and oxidizer elements arranged in concentric rings, the outermost ring providing fuel 
film cooling to the chamber. The chamber geometry is shown in Figure 12. The 
characteristic parameters of the injector are summarized in Table 1. 

As illustrated in Figure 13, the rig representing the Strutjet engine was designed as a 
sandwich with hinged side wall sections. The duct section housing the strutrocket had a 
fixed geometry of 4.0 in by 6.6 in. The isolator section in front of the strut duct could be 
connected to either a bell mouth for a static test or a hydrogen fueled vitiated air heater. 
Two struts were mounted in the strut duct, dividing the flowpath to the inlet into three 
channels. 

In the sea level static ducted rocket tests the isolator section in front of the strut duct was 
fitted with a calibrated bell mouth. The duct geometry was varied to determine the 
configuration yielding the maximum thrust. Sensitivities of rocket chamber pressure, 
mixture ratio, and rocket nozzle expansion ratio were also established. As shown in 
Figure 14, thrust enhancement was a strong function of the ram burner throat area and a 
somewhat weaker function of the rambumer geometry. With a duct geometry of 3°-3°, 
13% more thrust was obtained than for the reference rocket in a particular test with a 
throat area of 32 in 2 . Data analysis indicated that oxygen content of the inducted air was 
completely consumed in approximately 8 in. from the rocket baseline. Considering the 
influence of chamber pressure, operation at 2000 psia generated more thrust than 
operation at 1600 psia; however, airflow and thrust augmentation were reduced by 19% 
and 3%, respectively. The area ratio of 1 1:1 generated 12% higher induced air flow and 


13 



6% more thrust than the lower area ratio of 5:1. Finally, it was observed that greater air 
flow and higher thrust result from operation at higher mixture ratios due to reduced 
thermal choking resulting from the afterburning scheme. 

In the direct-connect ducted rocket tests which allow evaluation of the engine under flight 
trajectory conditions, the isolator section in front of the strut duct was connected to a 
hydrogen vitiated air heater with a Mach 2 nozzle. The duct geometry which, in the static 
tests, provided the maximum take-off thrust augmentation was maintained in these direct- 
connect tests. Measuring the duct pressure and assuming a particular inlet performance, 
allows for the determination of flight altitude and Mach number simulated in a given test. 
Figure 15 shows the thrust obtained in the ducted rocket and ramjet tests. The left branch 
of the figure depicts the thrust of the ducted rocket without additional fuel injection, and 
the right branch, that of the ramjet without rocket operation. The ducted rocket tests were 
conducted under fuel-rich conditions, the excess fuel sufficient to support 1 0 lbm/sec of 
air flow. The simulated trajectory provides 10 lbm/sec of air at approximately Mach 1 .5. 
Tests beyond Mach 1.5 were thus "lean" on an overall engine stoichiometric basis. 
Auxiliary fuel injectors could be used to increase the engine thrust. The peak thrust is 
seen to occur at a simulated altitude of 23,000 ft and Mach 2.85, with 31 lbm/sec of air 
being supplied to the engine. The peak thrust is over twice the bare rocket value, 
representing better than 100% thrust augmentation. 

Direct Connect Ram qnd_ Scramjet Tests 

The ramjet tests were conducted at Mach numbers of 2 and higher without rocket 
operation to optimize ramjet injector performance. The primary ramjet test variable, 
other than the injector parameters, was the fuel of choice. In support of the strategy for 
minimizing heat load and hot spots, the main emphasis was on achieving a short 
combustor length. 

In the scramjet operating regime three test series were conducted: 

In the first tests the scramjet geometry was explored and high combustion efficiency with 
a fixed geometry over the Mach number range of 2 to 8 were demonstrated. Mach 2 and 


14 




4 tests were conducted with JP-10 fuel. The Mach 8 tests simulated the effect of 
regeneratively heated fuel by using ethane instead of JP-10. 

In the second test series, the tests conducted with ethane were repeated with JP-10 fuel, 
using a slightly modified duct geometry, namely the first 12 in. downstream being of 
constant area followed by a 2° double-sided expansion over the remaining duct length. 
Auto-ignition was not achieved; however, when pilots were utilized for ignition and 
flame sustaining, stable combustion at 95% efficiency was observed. As shown in Figure 
16, the test duplicated, in essence, the performance previously achieved with ethane, with 
only the slight change in duct geometry. The Strutjet design used in these tests provided 
for a contact pilot at each injection point of the hydrocarbon ramjet fuel. This pilot 
derives its energy from the combustion of small amounts of the gelled rocket propellants 
which are injected and burned upstream of the hydrocarbon injection. Due to the 
hypergolic nature of the employed rocket propellants the pilots act initially as igniters and 
subsequently as flame sustainers, allowing flight at high Mach numbers and high 
altitudes. The demonstration of this feature is verified by the data presented in Figure 17. 
At a simulated flight condition of Mach 4 and 40,000 fit of altitude JP-10 was ignited by 
the pilot resulting immediately in a thrust increase of about 2,000 lbf. Combustion and 
thrust production were sustained as long as the pilot stayed on. When turned off, the 
combustion ceases and thrust collapses. 

The third test series was in support of the first freejet tests of the Strutjet engine to be 
conducted by NASA Lewis Research Center at their Hypersonic Test Facility (HTF) in 
Plum Brook. This facility has the capability to run simulated freejet flight conditions at 
Mach numbers of 5, 6 and 7 with a dynamic pressure of 1,000 psf. All previous strutjet 
direct-connect tests, were conducted at a dynamic pressure of 2,000 psf or higher. In 
addition the Strutjet testing had only been conducted at simulated flight conditions of M 
= 0-4 and 8. For the sake of risk reduction, additional direct-connect tests were 
conducted at Mach 6 and 7 at the reduced dynamic pressure. Figure 1 8 shows the test 
configuration and also the duct pressures achieved. These tests used the initial combustor 
divergence found efficient in the ducted rocket/ramjet test series. By properly staging the 
pilot and the unheated liquid JP-10 injection, good combustion efficiency was achieved 


15 



without reducing the duct divergence. This was a significant accomplishment since the 
employed engine geometry proved to be satisfactory for operation from Mach 0 to 8. 
Freeiet Tests 

The freejet engine shown in Figure 19 was designed constructed by Aerojet, delivered to 
and tested at Plum Brook. Tests were conducted with the identical fuel injection strategy 
used in the Mach 7 direct connect tests discussed above. These tests demonstrated inlet 
starting, fueled unstart with a large forward fueling split followed by inlet restart with 
shifting the fuel aft and substantial thrust increase. Figure 20 shows the internal pressure 
profiles and compares the freejet to the inlet data. Figure 21 shows the differential thrust 
produced as a function of the fuel flow. The slope break at an equivalence ratio of 0.55 is 
notable. This is the expected result of the non-uniform airflow distribution in the isolator. 
In the final test at HTF, the fuel injection distribution was shifted to better match the 
airflow. Unfortunately the facility experienced a hot isolation valve failure prematurely 
ending the test campaign. 

3.2 Gaseous Hydrogen System Tests 

All hydrogen systems considered for the Strutjet engine use cryogenic hydrogen. This 
hydrogen is used to cool the engine regeneratively. During this cooling process the 
hydrogen converts from a cryogen to a gas. All combustion related processes of the 
Strutjet engine will then use gaseous hydrogen as a fuel. Therefore, all combustion 
related tests were conducted with gaseous hydrogen. The tests described here were 
executed as part of the Advanced Reusable Technology program sponsored by NASA 
MSFC and supported by Aerojet with the objective to demonstrate the technology of the 
hydrogen fueled RBCC Strutjet engine previously described in section 2. 

Inlet Tests 

The previously shown Figure 4 illustrates schematically the air flow path on the vehicle 
underbody. Forebody compression reduces the inlet approach Mach number. For 
example, if the freestream Mach number is 6.0 then the approach Mach number is 
reduced to 4.2. A model of the inlet-isolator test article is shown in Figure 22. The inlet 
model accurately simulates the inlet from just upstream of the struts all the way through 
to the ram combustor. Geometrical similarity between the full-scale SSTO vehicle, 


16 



capable of delivering 25,000 Ibm payload to the International Space Station, and the 
subscale test article is maintained. The model is 6.8% scale of the full size inlet. The 
inlet is preceded by a plate simulating the vehicle forebody boundary layer. While the 
full-scale engine contains 8 to 16 struts and two sidewalls, the test article is composed of 
two struts and two sidewalls. In the model, the sidewalls are positioned to represent the 
symmetry plane between struts, thereby fully simulating flow around and between struts. 
The full-scale engine has two bleed locations, a forebody and a throat bleed. In order to 
adjust for the non-linear scale effects the inlet subscale model has an additional strut 
bleed which removes excess boundary layer build up on the sidewalls. A throat plug is 
used to simulate combustion pressure increase. 

The testing of this inlet in the NASA LeRC supersonic wind tunnel provided excellent 
results. Tests were performed over simulated flight Mach numbers from 3.6 to 8.1. The 
inlet started at all Mach numbers, and exhibited unstart margins of over 20% in ramjet 
and 100% in scramjet modes. With forebody spill excluded, engine capture efficiency 
exceeded 90%, which is remarkable, considering the wide operating range and large 
unstart margin realized with this inlet. The inlet started easily at all Mach numbers and 
generated excellent pressure rise, as shown in Figure 23. 

Strut Rocket Tests 

Aerojet and NASA Marshall Space Flight Center completed proof-of-concept testing on a 
new strut rocket injector element developed to enable operations under the unique 
strutrocket conditions. This element has been incorporated into the design of six subscale 
strut rockets for ducted rocket testing. 

The primary objective for a part of these tests was to demonstrate the performance and 
durability of a new injector which had been designed to maximize thruster efficiency 
while minimizing thruster length. The injector design ensures efficient mixing of the 
propellants within the chamber by employing an impinging element design and by 
utilizing a refined element pattern: 1 8 elements on the 0.5 inch diameter injector face. 
Figure 24 shows the injector firing in a single chamber in tests at Aerojet. Thermocouple 
data verified that the temperature of the injector face was within the limits predicted for 
the design. Post-test visual inspections of the test articles indicate that the injectors 


17 



suffered virtually no erosive or other damage related to excessive face temperatures. 
Scanning electron microscope images of single elements and single orifices produced at 
MSFC are particularly encouraging. The data obtained during 16 hot-fire tests extend 
over the ranges indicated in Table 2: 

Subscale strutrockets were fabricated and check out tests were complete. In these tests 
the propellant, coolant and ignition sequences required for RBCC operation were 
developed, and unaugmented rocket thrust was determined to establish a reference for 
subsequent combined cycle testing. Aerojet is also currently fabricating a full scale 
strutrocket using the same injector element. During the testing of this test article laser 
diagnostics will be used to measure the fuel distribution in the rocket exhaust for various 
design and operating conditions. 

Direct Connect Ram & Scramiet Tests 

Aerojet successfully demonstrated the efficiency of this fuel injection strategy at Mach 6 
and 8. Figure 25 shows one of the test strut assembly for the direct connect campaign. 
Two struts are shown installed in the test rig in Figure 26. 

Three test series were conducted: ram and scram tests at simulated Mach 6 flight 
conditions, and scram mode tests at Mach 8 conditions. High performance was achieved 
in all three test series in only 27 tests. A key to this success was the use of the cascade 
scram injectors in the forward location. These low drag supersonic injectors are shown in 
Figure 27. The fuel split for each mode and Mach number was determined by the inlet 
tolerance to the combustion pressure rise. Excellent performance at stoichiometric 
conditions was demonstrated at each Mach number and mode tested, by simple 
adjustment of the fuel flow to each of the three injectors. The data indicated that the 
performance of the cascade injector was even better than expected. Evidence of over- 
penetration suggested additional improvement can be made by increasing the number of 
cascades from 4 per strut side to 5 or more. This can be expected to increase the 
combustion rate and permit even shorter, lighter engine designs. 

Planned Tests 


18 




In the first two tests strutrockets, shown in Figure 28, will be installed in the direct 
connect duct to explore the performance in the Scram/Rocket and Ascent/Rocket Modes. 
A second pair of strutrockets will be installed in the new ffeejet engine and tested first 
under sea level static conditions and then in the ducted rocket and ramjet modes. Of 
particular interest is the transition from ducted rocket to ramjet mode at a flight Mach 
number of about 2.5. Freejet test hardware is shown in Figures 29, 30, and 31, showing 
the inlet variable geometry, the integration of the strutrockets into the combustor, and the 
overall cross section of the flowpath. Comparison between Figures 4 and 3 1 reveal the 
operation of the inlet variable geometry. The nozzle flap shown in Figure 3 1 is integrated 
in the test hardware on the body side of the engine for test arrangement purposes; in the 
flight engine the nozzle flap will be on the opposing cowl side. 

The contemplated test facility has the capability to provide accelerating test conditions. 
This allows tests to begin at one Mach number and sweep continuously to a higher one 
while simultaneously matching pressure and enthalpy. Using this facility, it will be 
possible to demonstrate the ducted rocket to ramjet and the ram to scram transitions. 

4.0 Maturity of Required Strutjet Technologies 

The maturity of RBCC engines is best defined in terms of a ‘Technology Readyness 
Level” (TRL) as defined by NASA. The following assessment is as of Summer 1998 
relative to a TRL of 6 which requires technology demonstration in a relevant 
environment be it simulated on the ground or actually flown. Table 3 categorizes various 
aspects of an RBCC engine into the degree of maturity relative to TRL 6. 

It is evident from the table that a large amount of development must be accomplished 
before a TRL of 6 is achieved for the Strutjet RBCC engine. With this assessment in 
mind, it may not be prudent to risk at this point in time the commitment of large amount 
of resources towards the exclusive RBCC approach to achieve low cost access to space. 
For near term applications alternate approaches, like further maturation of all rocket 
propulsion, should be employed. However, the potential of RBCC engines is so 


19 



overwhelming that it is also not prudent to casually dismiss the opportunity to exploit the 
RBCC option. 

5.0 Description Of A Rapid Cargo Delivery System 

Mission Description - For the rapid cargo delivery systems (RCDS) considered, it is 
assumed that the concept of a multy hub concept is employed. Between hubs, which 
typically would be separated by oceans, the cargo is hypersonically transported using the 
RCDS; to and from hubs, the cargo is flown via subsonic aircraft. If hubs are located near 
an ocean or in areas of low population density the hypersonic portion of the mission can 
be accomplished with ground launched RCDS. If these hub conditions are not given, air 
launched RCDS vehicles can be carried on top of large transport aircraft to 
environmentally feasible launch sites and then released. Since the entire concept is only 
economically more feasible if very high launch rates are considered, the environmental 
impact resulting from combustion exhaust gases and noise generation will have a very 
strong influence. In order to bring the operating cost down, hydrogen fuel has to be 
abondoned. The fuel of choice, and the one further pursued in this paper, is cryogenic 
propane. It is anticipated that this fuel will simplify the fuel supply infrastructure, and 
fuel loading operations. 

The baseline missions selected for cargo delivery are summarized in Table 4. Since the 
emphasis of this study is on Pacific rim markets two main Western United States based 
hubs are considered: Los Angeles and Anchorage. The primary destination hubs, shown 
in Figure 32, are in other Pacific rim countries, except for a hub in Kirma, Sweden, which 
can easily be reached via a polar route and could open up the European market. The 
baseline RCDS vehicle designs are sized to cover the maximum range needed; shorter 
range hubs can be accommodated using the same vehicles through propellant off loading. 
The maximum ranges to be achievable in this mission model are about 7620 mi or 
14,1 1 1 km, one being Los Angeles to Singapore, and the other, with an equivalent range, 
Anchorage to Perth, Australia. 


20 



RBCC Engine Model - The baseline engine model used in the following vehicle designs 
and trajectory simulations uses liquid oxygen (L0 2 ) and liquid propane (LC 2 H 8 ) with a 
mixture ratio of 3. This engine is in principle very similar to the engine previously shown 
in Figures 2, 3 and 4. A qualitative comparison of the predicted performance of the 
hydrogen and the propane Strutjets in terms of specific impulse is given in Figure 33. 
Again, the ducted rocket mode provides some enhanced impulse over the pure rocket 
mode up to Mach 2.5, where the ramjet mode initiates. The scramjet mode carries out to 
a cutoff at the transition Mach (M T ) number of 10. The specific impulse is the “cowl-to- 
tail” (no forebody drag) engine performance. In the simulations, the effective Isp is 
reduced by the vehicle/engine atmosphere drag as shown in Figure 34. At the transition 
Mach number the engine will first operate in the combined scramjet/ascent rocket mode, 
and then gradually tranition into the ascent rocket mode for the remainder of the bum. 
After burnout the Strutjet engine is shut down and not used anymore for the remainder of 
the mission because the vehicle will enter a glide phase and land at the target hub 
unpowered.. 

The RBCC engine thrust-time history for the ground-launched vehicle is shown in Figure 
35. This thrust profile addresses all modes of power-on operation. The selected capture 
area of the installed Strutjet engine is 255 ft 2 , and during the airbreather mode of 
operation the dynamic pressure is assumed to be 2,000 psf. The ducted rocket mode is 
terminated at an altitude of about 30 kft. The ascent-rocket or pure-rocket mode is 
initiated at an altitude of 98 kft, and burnout occurs at an altitude of 1 80 kft. 

Ground-Launched Vehicle Design Baseline - While an earlier paper 4 described the 
L0 2 /LH 2 RBCC orbital vehicle design trades for carrying 25 klbm to the International 
Space Station orbit, this paper contains a baseline L0 2 / LC 2 H 8 vehicle design of a ground 
launched (horizontal takeoff/horizontal landing) rapid cargo delivery system (RCDS-GL) 
sized to deliver 5 klbm to various bases, primarily to Pacific rim locations. This baseline 
vehicle is compared to the Space Shuttle and the previously established orbital RBCC 
vehicle in Figure 36, and shown in more detail in Figure 37. The gross liftoff mass of 
this RCDS-GL is about 477 klbm, and the dry mass is estimated to be 79 klbm; usable 


21 



propellant is about 373 klbm as shown in the table contained in Figure 37. This vehicle 
design provides 29 klbm for runway propellant, which leaves the ignition mass at 506 
klbm. This vehicle is about 1 10 ft long, has a wing span of 50 ft, and requires two RBCC 
engines each with 143 klb thrust to achieve a vehicle lift-off thrust-to- weight ratio of 
0.6:1. The cargo bay of 10 x 20 ft is compatible with a typical 5,000 lbm cargo volume. 
The RBCC engine thrust-to- weight is assumed to be 43:1, which due to the use of the 
higher density propane fuel is substantially higher than that for the hydrogen engine. The 
vehicle total velocity is sized to achieve 26,900 fps, which includes an axial maneuvering 
velocity reserve of 1,000 fps for landing. The reaction control system (RCS) velocity of 
200 fps is sufficient for exoatmospheric attitude control while gliding down from 200 kft 
to about 50 kft where aerodynamic surfaces are able to control the landing maneuvers 
required. 

Ground-Launched Vehicle Design Trades - Several ground-launched RBCC vehicle 
designs were evaluated. The vehicle ignition mass is shown in Figure 38 as a function of 
vehicle thrust-to-weight ratio at ignition ranging from 0.6 to 0.8 for transition Mach 
numbers of 10 and 12, and propane RBCC engine thrust-to-weight ratios ranging from 25 
to 43:1. Two hydrogen RBCC cases are presented for comparison purposes, showing that 
a lighter vehicle (by about 30 to 40%) is possible when using hydrogen propellant in lieu 
of propane. The vehicle mass is seen to increase as greater ignition thrust is needed. 
However, the baseline vehicle design chosen for this trajectory analysis is shown at an 
ignition weight of 506 k lbm, an engine T/W of 43:1, and an M x of 10. 

Ground-Launched Trajectory - The baseline RCDS-GL vehicle is sized to accommodate 
the maximum range missions from two US hubs to cover the mission model as previously 
indicated in Table 4. The flight of a ground-launched vehicle is simulated on the Boeing 
OTIS trajectory simulation code which has the attribute to maximize range. The flyout 
trajectory to burnout is presented in Figure 39 for this typical vehicle design. The liftoff 
velocity is 450 fps at an altitude of 200 f., and an angle of attack of 17.2 deg for 
appropriate subsonic lift conditions. The propellant weight used on the runway is 29 
klbm. The vehicle flies up an altitude-range profile at Q=2,000 psf while achieving the 
transition Mach number of 1 0 at a flight time of about 400 sec. Then, the ascent-rocket 


22 



mode takes over and the vehicle flies up to an altitude of 179 k ft which is reached at 
burnout after a flight time of 610.3 sec. 

A typical maximum range trajectory of 7,643 nmi or 14,154 km (from Anchorage, 
Alaska to Perth, Australia) is presented in Figure 40. At rocket burnout the velocity 
achieved is 20,523 fps at a surface range of 850 nmi. The vehicle then coasts up to 190 
kft before a long glide down into the denser atmosphere, achieving Mach 2 after a flight 
time of 4,000 sec (67 min) at an altitude of 60 kft. The vehicle’s altitude descent 
continues, achieving Mach 1 at 20 kft, and subsonic speed at 5 kft after a flight time of 
4,400 sec (73 min.). The axial rocket thrusters are available to maneuver the vehicle to 
the final target area while gliding to a final unpowered landing. 

The vehicle flies a high dynamic pressure (Q) profile during the airbreathing mode to 
Mach 10, whereupon the ascent-rocket mode takes over to achieve a Mach number of 
about 19 at a low Q of about 200 psf, as seen in Figure 41. As the vehicle coasts up and 
passes over the apogee to glide back to lower altitude, the Q builds up to reach about 
1,000 psf at 100 kft; as the vehicle descends to 40 kft, the Q drops to about 200 psf. 
Transitioning the denser atmosphere at near sonic speeds the vehicle is subjected to a 
higher Q (600 psf) prior to dropping off at 5,000 ft altitude to under 100 psf. 

Several ground-launched trajectory cases have been evaluated for maximum vehicle 
temperature and heat flux experienced during the mission. The baseline heat flux is 200 
BTU/ft 2 -sec, which corresponds to a stagnation temperature of 3,750 deg-F, as presented 
in Figure 42. As seen, the heating environment constraint decreases as the maximum 
range achievable drops off. A range decrease of 28 percent results for a maximum heat 
flux of 100 BTU/ft 2 -sec with a temperature of 3,100 deg-F on the nose. The thermal 
protection systems previously postulated for an orbital vehicle used carbon/silicon 
carbide (C/SiC) on the nose and leading edges of the vehicle, graphite-epoxy for the body 
structure and LH 2 tanks, and aluminum-lithium for the L0 2 tanks. At the higher heat flux 
levels of the RCDS-GL the C/SiC may have to be replaced with hafnium diboride (HfB 2 ) 
on the nose and leading edges. 

The heat flux and nose stagnation temperatures for the baseline vehicle as a function of 
Mach number are shown, respectively, in Figures 43 and 44. The heat flux ascent profile 


23 



reaches 200 BTU/ft 2 -sec near burnout and stays at this constraint during the initial glide 
portion of the descent trajectory from apogee. As the vehicle reenters the atmosphere the 
heat flux and stagnation temperature drop off; the latter follows the ascent heating 
temperature closely to the target landing area. 

The ground-launched vehicle mission propellant mass requirements are 
summarized in Figure 45. This design trade is for a Los Angeles to Tokyo mission with 
an engine thrust-to-weight ratio of 35:1 and a maximum heat flux of 150 BTU/ft 2 -sec. 
The baseline vehicle in this case is flown to a maximum range of 6,793 nmi (13,580 km). 
If the heat flux is constrained to 100 BTU/ft 2 -sec then the maximum range is limited to 
only 5,515 nmi (10,213 km), or a loss of about 20 percent. Off-loading propellant from 
the baseline vehicle results in shorter ranges achievable. The baseline useable propellant 
mass is 373 klbm for the maximum range, which carries the vehicle 605 nmi beyond 
Brisbane from Los Angeles to near Perth. Off-loading propellant to 250 klbm is required 
to reach Tokyo from Los Angeles. A short hop to Anchorage from Los Angeles requires 
the loading of only 90 klb of propellant. 

Air-Launched Vehicle Design Baseline - An air-launched vehicle (RCDS-AL), depicted 
in Figure 45, is also sized to cover the same mission model. The vehicle selected is sized 
in this case for an engine thrust-to-weight ratio of 35:1 and a vehicle thrust-to-weight 
ratio at ignition of 1.0 after aircraft separation. The total velocity used in the vehicle 
sizing is 23,900 fps, plus the same onboard maneuvering capability as for the ground- 
launched design. The gross takeoff mass of this vehicle is 390 klbm, while the useable 
propellant is 290 klbm. The vehicle layout, shown in Figure 46, is similar to the ground- 
launched vehicle previously shown in Figure 37. Depending on the propellant tank sizes 
the air-launched vehicle could be shorter, as the tank volumes are less for the air-launched 
design. 

The launch platform chosen for this air-launched design is the Russian AN-225, as 
depicted in Figure 47. This aircraft has a payload limit of about 550 klb. The launch 
conditions assumed the RCDS-AL vehicle is released at an altitude of 40,000 ft and a 


24 



Mach number of 0.8 (775 fps). After separation the RBCC engines ignite and start 
operation in the ducted rocket modes. 

Air-Launched Vehicle Design Trade - Several air-launched vehicle designs have been 
evaluated. The vehicle gross ignition mass is shown in Figure 48 as a function of vehicle 
thrust-to-weight ratios ranging from 0.6 to 1 .0, for transition Mach numbers of 10 and 12, 
and propane Strutjet engine thrust-to-weight ratios of 30, 35 and 43:1. The vehicle mass 
decreases for the lower thrust levels, but the maximum range is achieved at the baseline 
design point with a thrust-to- vehicle weight ratio of 1.0. Again the baseline OTIS 
trajectory simulation is shown for an engine T/W of 35:1 and an M T of 10, resulting in a 
vehicle gross weight of 390 klbm. 

Air-Launched Trajectory - The baseline vehicle is sized to fly a maximum range from 
both US hubs. The vehicle flight is simulated on the OTIS trajectory code from aircraft 
launch over the Pacific Ocean to the target hub. The baseline vehicle flies up the high 
dynamic pressure (Q = 2,000 psf) profile to about 100 kft altitude, then the ascent-rocket 
mode carries the vehicle to a burnout at 181 kft altitude after a flight time of 498 sec. (8.3 
min), as depicted in Figure 49. The vehicle’s burnout velocity is 20,583 fps at a Mach 
number of about 19.5. 

Figure 50 presents the altitude-range profile of the baseline vehicle flown from 
Anchorage to the maximum range. The baseline vehicle is able to cover Singapore and 
Brisbane with sufficient range to land near Perth, Australia. The flight profile is similar 
to the ground-launched vehicle, previously shown in Figure 40. The maximum range 
achieved is 7,856 nmi (14,548 km), and the vehicle reaches an apogee altitude of about 
200 kft prior to the long glide back to the lower atmosphere. The dynamic pressure 
buildup is also similar to the ground-launched case, and reentry at Mach 1 (20 kft 
altitude) is achieved after 4,200 sec (70 min) flight time. Another 100 sec is taken to 
reach the target area flying subsonically. The onboard reaction-control system (RCS) is 
used to stabilize the vehicle during the glide phase from apogee. The onboard 
maneuvering rocket system (OMS) of 1 ,000 fps velocity capability allows the vehicle to 
fly to alternate bases and/or readjust the range, as needed, during the flight. 


25 



The air-launched vehicle mission propellant mass requirements are summarized in 
Figure 51. This design trade example is for flights from the Anchorage base to shorter 
ranges. For flights to Brisbane the propellant loading needs to be only 80 percent of the 
nominal 290 klbm propellant. For flights to Tokyo the loading reduces to 39 percent of 
maximum. Off-loading propellant from the baseline vehicle results in the shorter ranges 
achievable. 

7.0 Summary and Conclusions 

Two propane fueled rapid cargo delivering systems vehicles have been presented capable 
to operate at hypersonic speeds between hubs around the Pacific rim countries. With 
flight times between hubs below 90 min, and the ability for either ground or airlaunch, an 
attractive new transportation system can be envisioned. This system is suitable for 
performing its missions in an environmentally acceptable fashion relative to the 
generation of noise and air pollution. Noise is mitigated by subsonic shuttle service of the 
cargo to and from the hubs, and air launch of the hypersonic vehicle to an noise 
insensitive launch site. Air pollution is minimized through the use of relatively clean 
burning propane propellant. These vehicles may have the potential to operate eventually 
in an airline type fashion with long mean times between overhaul, bringing operating cost 
down levels where new markets will open up. The key to this potential is the 
abandonment of the power density operation of all-rocket propulsion in favor of a more 
benign rocket based combined cycle propulsion concept. 

Hydrogen And Hydrocarbon Strutjet Engines - These engines, particular the hydrogen 
fueled one, have made considerable progress, and test data established to date verify their 
fundamental feasibility and support earlier performance predictions. While the current 
NASA Advanced Reusable Technology (ART) Program as well as Aerojet sponsored test 
activities will provide additional data during 1998 and early 1999, no further design or 
test activities are currently planned for storable hydrocarbon RBCC engines. 


26 



Struiet Technology Maturity - Although significant achievements have been made 
towards a Technology readiness Level of 6, demonstration in a relevant environment, 
most of the effort to date is focused on flowpath development and performance 
assessment. Future work must be done on engine structure, thermal management, and 
propellant feed system. Ground tests of flight type engines are mandatory before 
committing these advanced engines to in-flight, captive carry or self-powered, evaluation. 

Rapid Cargo Delivery System - Advanced RBCC Strutjet vehicle designs using 
L0 2 /LC 2 H 8 propellants have the potential to offer robust vehicles capable to perform 
rapid cargo delivery missions to various points around the Pacific rim to and from US 
hubs. Even though a L0 2 /LH 2 design would have a lower gross weight, the use of 
propane in lieu of hydrogen allows easier handling and storage of propellants. Air- 
launched vehicles flying from an AN-225 aircraft may be safer with propane onboard 
than with a hydrogen vehicle. Vehicles transporting a 5,000 lb cargo may offer delivery 
of the payload in flight times of about 1 hour, depending on the mission base destination. 
Ground vehicles weighing about 480 klbm and air vehicles weighing about 390 klbm 
could be designed to reach the major city ports around the Pacific rim, and possibly to 
Europe when flying over the North Pole. 

Advanced thermal protection system (TPS) materials may have to be developed to handle 
the heat flux and leading edge temperature environments experienced in flying ranges out 
to 14,000 km. Maximum range achievable is decreased for lower heating design 
constraints. 

The vehicle/engine designs presented establish first order feasibility of rapid cargo 
delivery systems. If future economic studies can support the need and profitability of such 
system, more in depth engine and vehicle designs are required and recommended. 

7.0 References 

1. William Scott, “ ” Aviation Week, June 15, 1998 

2. Penn J.A. and Lindley C.A., The Aerospace Corporation, 2350 E. El Segundo Bl., El 
Segundo, California, 90245-4691, “Requirements And Approach For A Space 
Tourism Launch System” 


27 



3. Bulman, M.J. and Siebenhaar, A., Aerojet Propulsion Division, Sacramento, 
California, “ Rocket Based Combined Cycle Propulsion For Space Launch”, 46th 
International Astronautical Congress, 1995 / Oslo, Norway, IAF-95-S.5.02 

4. Siebenhaar, A., GenCorp Aerojet, Sacramento, California, and D.K. Bonnar, The 
Boeing Phantom Works, Huntington Beach, California, “Strutjet Engine Paves Road 
To Low Cost Space Access”, 48th International Astronautical Congress, 1997/ Turin, 
Italy, IAF-97-S.5.05 

5. US Patent No. 5,220,787 “ Scramjet Injector”, issued to Aerojet June 22, 1993 


28 



Table 1: 


Parameter 

Design 

Nominal Operation 

Mixture Ratio 

1.6 

1.4 

Chamber Pressure (psia) 

2,500 

1600-2000 

Thrust (lbf) 

1,000 

600-700 

Expansion Ratio 

5.1 and 11.1 

5.1 and 11.1 


















Table 2: 











Table 3: 


Top Level Technology Item 


Flowpath Integration 


Fuel Injection 


Light Weight Structure 


Strut Rockets 


Turbo Pumps 


Engine Controls 


Testing 


Sub Tier Technology Item 


Design Point Selection 


Balanced Performance Along 
Flowpath 


Penetration, Mixing, Vaporization 


Stable Combustion 


Controlled Heat release 


High Temperature Materials 


Radiation Cooled Structures 


Regeneratively Cooled Structures 


Endothermic Fuel Reactions 


Closed Loop Cooling 


High Thrust-To-Mass 


Long Life 


Low Weight 


Long Life Bearings 


Low Temperature Turbines 


Multi Mode Operation 


For All Flight Modes 


Ground Test Subscale 


Ground test Full Scale 


Captive Flight Of Engine Module 


Self-Powered Flight Of Flight Type 
Engine 


Degree of Technology 
Maturity 


Low I Medium | High 
















































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